Користувач:Naelsia/Ariane-1

Джерела:

1 ESA Bulletin №1 June 1975 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

2 ESA Bulletin №3 October 1975 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

3 ESA Bulletin №4 February 1976 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

4 ESA Bulletin №5 May 1976 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

5 ESA Bulletin №6 August 1976 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

6 ESA Bulletin №8 February 1977 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

7 ESA Bulletin №9 May 1977 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

8 ESA Bulletin №11 October 1977 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

9 ESA Bulletin №12 February 1978 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

10 ESA Bulletin №13 May 1978 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

11 ESA Bulletin №15 August 1978 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands. (головне)

12 ESA Bulletin №16 November 1978 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

13 ESA Bulletin №17 February 1979 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

14 ESA Bulletin №18 May 1979 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

15 ESA Bulletin №19 August 1979 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

16 ESA Bulletin №20 November 1979 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

17 ESA Bulletin №21 February 1980 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

18 ESA Bulletin №22 May 1980 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

19 ESA Bulletin №23 August 1980 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

20 ESA Bulletin №24 November 1980 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

21 ESA Bulletin №25 February 1981 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

22 ESA Bulletin №26 May 1981 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

23 ESA Bulletin №27 August 1981 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

24 ESA Bulletin №28 November 1981 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

25 ESA Bulletin №29 February 1982 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

26 ESA Bulletin №30 May 1982 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

27 ESA Bulletin №31 August 1982 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

28 ESA Bulletin №32 November 1982 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

29 ESA Bulletin №33 February 1982 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

30 ESA Bulletin №34 May 1983 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

31 ESA Bulletin №35 August 1983 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

32 ESA Bulletin №36 November 1983 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

33 ESA Bulletin №37 February 1983 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

34 ESA Bulletin №38 May 1984 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

35 ESA Bulletin №39 August 1984 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

36 ESA Bulletin №43 August 1985 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

37 ESA Bulletin №45 February 1986 - ESA Scientific and Technical Information Branch, clo ESTEC, Oomeinweg, NOOROWIJK, Netherlands.

38 https://www.esa.int/About_Us/ESA_history/Antonio_Fabrizi_From_nuts_and_bolts_to_Europe_s_launchers_of_tomorrow

Info

ARIANE

Participating ESA Member States: Belgium, Denmark, France, Germany, Italy, Netherlands, Spain, Sweden, Switzerland Planned first test flight: - Mid-1979

The purpose of the programme is to enable Europe, as from about 1980, to place geostationary satellites in orbit. The launcher is being designed to place payloads of about 1500 kg in transfer orbit, which in turn will enable satellites of the orde r of 750 kg to be injected into geostationary orbits by means of a suitable apogee motor. ARIANE is a three-stage vehicle, 47.6 m high and weighing 202 tonnes at lift-off. The first stage, L 140, with a diameter of 3.8 m, contains 140 tonnes of propellants (N2 0 4 and UDMH) and is powered by four Viking-2 engines, with turbo-pump and film-cooled single-wall nozzle. The total thrust at lift- off is 240 tonnes and the burn time 150 s. The second stage, L 33, with a diameter of 2.6 m, carries 33 tonnes of the same propellant and is equipped with a Viking-4 engine derived from the Viking-2 by adapting the nozzle for operation in vacuum. The th ird stage, H 8 (same diameter as L 33), carries 8 tonnes of liquid hydrogen and liquid oxygen and is powered by an HM 7 engine with a thrust of 6 tonnes. Stage separation is achieved by explosive cords, and the stages are forced apart by retro and acceleration rockets. The intrument pack, located above the third stage, performs the autopilot, guidance and sequencing functions centrally with the help of a computer. It also carries telemetry, telecommand trajectography and break-up equipment, together with the inertial platform. The fairing is bulb-shaped and can house a spacecraft payload 3 m in diameter and 4 m high. It is planned to qualify the launcher by means of four test flights, starting in mid-1979, in order to achieve a fully operational launcher by the beginning of the 1980s. The execution of the programme is entrusted to the Centre National d'Etudes Spatiales (CNES)' control of the execution of the programme being exercised by ESA.

АРІАН

Участь країн-членів Європейського космічного агентства: Бельгія, Данія, Франція, Німеччина, Італія, Нідерланди, Іспанія, Швеція, Швейцарія Перший запланований випробувальний польот: - середина 1979 року

Мета програми - забезпечити Європу можливість запускати геостаціонарні супутники на орбіту від близько 1980 року. Ракета-носій розробляється для запуску корисних навантажень вагою приблизно 1500 кг на трансферну орбіту, що в свою чергу дозволить впроваджувати супутники вагою приблизно 750 кг на геостаціонарні орбіти за допомогою відповідного апогейного двигуна. АРІАН - це триступінчастий апарат заввишки 47,6 м та масою 202 тонни при старті. Перший ступінь, L 140, з діаметром 3,8 м, містить 140 тонн палива (N2O4 та UDMH) і працює на чотирьох двигунах Viking-2 з турбонасосом та охолоджуваним плівковим одношаровим соплом. Загальний тяговий потік при старті становить 240 тонн, а час горіння - 150 с. Другий ступінь, L 33, з діаметром 2,6 м, містить 33 тонни того ж палива і обладнаний двигуном Viking-4, що походить від Viking-2, з адаптованим соплом для роботи в вакуумі. Третій ступінь, Н 8 (такого ж діаметру, як L 33), містить 8 тонн рідкого водню та рідкого кисню і працює на двигуні HM 7 з тягою 6 тонн. Відокремлення ступенів досягається за допомогою вибухових шнурів, а ступені розділяються за допомогою ретро- та прискорюючих ракет. Приладовий блок, розташований над третім ступенем, виконує функції автопілота, навігації та послідовності централізовано за допомогою комп'ютера. Він також несе телеметрію, телекоманди, траєкторію руху та обладнання для розпаду, разом з інерційною платформою. Обтічник має форму кулі і може вміщувати космічний корисний завантаження з діаметром 3 м та висотою 4 м. Планується пройти кваліфікацію ракети шляхом проведення чотирьох випробувальних польотів, починаючи з середини 1979 року, для досягнення повноцінної робочої ракети на початку 1980-х років. Виконання програми доручено Національному центру космічних досліджень (CNES), контроль за виконанням програми здійснюється Європейським космічним агентством (ESA).(1)

First firing of the complete third-stage HM7 engine A complete HM7 engine was fired by SEP on the horizontal test stand at Villaroche for the first time on 27 May 1975. The test lasted 10 s, including 5 s at full thrust, and took place without a hitch. It was the first of a series. of tests intended to finalise the start-up sequence for the complete engine. The soundness of the start-up principle as well as the correct operation of the complete engine and the test stand were demonstrated at this first test. The HM7 engine powers the launcher's third stage (H8). It burns liquid hydrogen and liquid oxygen and produces a thrust of 60 kN for a chamber pressure of 30 bar. SEP has entrusted the development and manufacture of the combustion chamber's injection head and of the nozzle to MBB, Munich, while the development and manufacture of the turbopump and the integration and testing of the complete engine are being carried out by SEP itself at Vernon. Passenger experiments aboard Ariane test flights The proposal by ESA in March 1975 to fly passenger experiments on Ariane test flights aroused considerable interest. By 18 June, 77 experiment proposals had been received, and further propositions continue to arrive. All types of experiments are represented (scientific, technological and application), two of the proposals emanating from countries outside the Member States of ESA. A briefing meeting organised by ESA at Frascati on 19 and 20 June brought together some 50 of the potential experimenters. The first day was devoted to a presentation of the launcher and the Ariane programme, and the second to a review of the experiment propositions in which interest had been expressed. The review revealed several proposals relating to apogee motors and transfer modules. If these were to be flown, the range of experiments that could be achieved would be enlarged since, in addition to the 200/36000 km elliptical orbit currently available, geosynchronous orbit could be envisaged for launches L03 (mid 1980) and/or L04 (end 1980). It is expected that firm propositions will have been received by 30 September 1975. Thereafter, ESA will carry out the selection phase with a view to arriving at decisions by mid-1976. Following the Frascati meeting, consultation with the various potential experimenters is in progress to establish a first approach to possible configurations and their associated integration problems

Перше запуск повного двигуна третьої ступеня HM7 Повний двигун HM7 був запущений SEP на горизонтальній випробувальній установці у Вільярош вперше 27 травня 1975 року. Випробування тривало 10 секунд, включаючи 5 секунд на повну тягу, і пройшло без проблем. Це було першим з серії випробувань, спрямованих на узгодження послідовності запуску для повного двигуна. На цьому першому випробуванні було продемонстровано правильність принципу запуску, а також правильну роботу повного двигуна та випробувальної установки. Двигун HM7 приводить у рух третій щабель ракети (H8). Він спалює рідкий водень та рідкий кисень і виробляє тягу 60 кН при тиску камери 30 бар. SEP доручив розробку та виробництво вприскувальної головки камери згорання та сопла компанії MBB, Мюнхен, тоді як розробка та виробництво турбонасоса та інтеграція та випробування повного двигуна здійснюються самою SEP у Верноні. Пасажирські експерименти на борту тестових польотів Аріана Пропозиція ЄКА в березні 1975 року про проведення пасажирських експериментів на тестових польотах Аріана викликала значний інтерес. К 18 червня було отримано 77 пропозицій експериментів, і надалі продовжували надходити інші пропозиції. Представлені всі види експериментів (наукові, технологічні та практичні), дві з пропозицій надійшли від країн поза країнами-членами ЄКА. Зустріч-консультація, організована ЄКА у Фраскаті 19 і 20 червня, об'єднала близько 50 потенційних експериментаторів. Перший день був присвячений презентації ракети та програми Аріана, а другий - огляду пропозицій експериментів, щодо яких був виявлений інтерес. Огляд показав кілька пропозицій, пов'язаних з апогейними двигунами та модулями передачі. Якщо їх буде запущено, діапазон експериментів, які можна буде здійснити, буде розширено, оскільки, крім еліптичної орбіти 200/36000 км, яка в даний момент доступна, можливо передбачити геосинхронну орбіту для запусків L03 (середина 1980-х) і/або L04 (кінець 1980-х). Очікується, що до 30 вересня 1975 року будуть отримані конкретні пропозиції. Після цього ЄКА проведе етап відбору з метою прийняття рішень до середини 1976 року. Після зустрічі в Фраскаті триває консультація з різними потенційними експериментаторами для встановлення першого підходу до можливих конфігурацій та пов'язаних з ними проблем інтеграції.(2)

VISIT TO THE ARIANE FACILITIES AT LES MUREAUX AND VERNON BY DELEGATIONS TO THE ARIANE LAUNCHER PROGRAMME BOARD On 9 December 1975, the delegations to the Ariane Launcher Programme Board visited the main facilities set up at Les Mureaux and Vernon under the Ariane programme. SNIAS (Les Mureaux) is responsible both for launcher integration (as industrial architect) and for developing the structures of the three stages and the fairing and integrating them. To this end, SNIAS is setting up the Launcher Integration Site (Site d'lntegration Lanceur - SIL), the buildings of which are complete and which is being fitted out according to schedule. The first operation to be carried out at the site will be the integration of the various stages intended for the dynamic mock-up of the launcher, on which tests will start in mid-1976. The visitors were also able to see the construction of the first-stage tank; three tanks have now been completed and two others are being manufactured. SEP (Vernon) is responsible for developing and manufacturing the propulsion systems for the three stages. At Vernon, the visitors witnessed an 80 s firing of a Viking-I I engine on test-stand PF2. In addition, they were able to note progress on the cryogenic test-stands PF41, 42 and 43, intended for the development and qualification of the third-stage propulsion system. Teststand PF41 (intended for turbo-pump and engine tests) and the control room are practically complete, PF41 being due to enter service in March 1976. The visitors also saw test-stand PF20, which is primarily intended for firing the complete first stage. After the installation and qualification of the control and measurement facilities, scheduled for Spring 1976, the first cluster fi"dng (thrust frame equipped with four Viking-I I engines) will take place there early in September 1976. It should be recalled that the tests on the propulsion system of the second stage L33 (engines and stage) will be carried out by DFVLR at Lampoldshausen in Germany. They will start in April 1976.

ВІЗИТ ДЕЛЕГАЦІЙ ДО ОБ'ЄКТІВ АРІАН В ЛЕ МЮРО ТА ВЕРНОН В РАМКАХ ЗАСІДАННЯ ПРОГРАМНОЇ РАДИ ПРОГРАМИ ЗАПУСКУ АРІАН 9 грудня 1975 року делегації Ради програми запуску Аріан відвідали основні об'єкти, створені в Ле Мюро та Вернон в рамках програми Аріан. SNIAS (Ле Мюро) відповідає як за інтеграцію ракети (як промисловий архітектор), так і за розробку конструкцій трьох ступенів, обтічника та їх інтеграцію. Для цього SNIAS створює Сайт Інтеграції Ракети (Site d'lntegration Lanceur - SIL), будівлі якого завершені, і який обладнується за графіком. Першою операцією, яка буде проведена на сайті, буде інтеграція різних ступенів, призначених для динамічної макетної ракети, на якій почнуться випробування наприкінці 1976 року. Відвідувачі також могли побачити будівництво бака першого ступеня; три баки вже завершені, а два інші виготовляються. SEP (Вернон) відповідає за розробку та виготовлення систем приводу для трьох ступенів. В Верноні відвідувачі дивилися 80-секундне випробування двигуна Viking-I I на випробувальному стенді PF2. Крім того, вони могли побачити прогрес на криогенних випробувальних стендах PF41, 42 та 43, призначених для розробки та кваліфікації системи приводу третього ступеня. Випробувальний стенд PF41 (призначений для випробувань турбонасосу та двигуна) та диспетчерська кімната практично завершені, очікується, що PF41 почне працювати в березні 1976 року. Відвідувачі також побачили випробувальний стенд PF20, який передбачений перш за все для запуску повного першого ступеня. Після встановлення та кваліфікації засобів контролю та вимірювання, запланованих на весну 1976 року, перший кластерний запуск (тягова рама з чотирма двигунами Viking-I I) відбудеться там на початку вересня 1976 року. Слід зазначити, що випробування системи приводу другого ступеня L33 (двигуни та ступінь) будуть проведені DFVLR в Лампольдсгаузені в Німеччині. Вони розпочнуться у квітні 1976 року. (3)

A 720 ton tower was raised by 6.5 m in seven hours at the Ariane Launch Site in Guiana on 24 March . The task was to raise the 50 m high Europa 11 tower to make it compatible with the Ariane launcher, which is 47 m high compared with the 32 m of the Europa II rocket. The tower's new lower section was erected during the next two days. Jointing work began immediately afterwards and is expected to be finished late in April. This relatively long time span is due to the rain, which makes welding particularly difficult. The remaining civil-engineering work on the launch site is on schedule. The launch table foundation is 60% complete and will be finished by the end of July . Work on the propellant-storage areas has started recently, and the situation is now as follows: 40 UDMH: retention lagoon protecting wall completed retention lagoon completed work started recently. Construction of the iced-water tank has also begun. This will constitute a reserve for cooling stored propellants and for airconditioning the premises within the launch table foundation and the Control Centre.The roadways have been laid out and will be complete at the end of the summer. The assembly building has been refurbished and the acceptance inspection should take place in May.

На стартовому майданчику "Аріан" в Гвіані підняли вежу вагою 720 тонн на 6,5 метра за сім годин 24 березня. Завданням було підняти 50-метрову вежу "Європа 11", щоб зробити її сумісною з ракетоносієм "Аріан", який має висоту 47 метрів порівняно з 32 метрами ракетоносія "Європа II". Новий нижній розділ вежі було встановлено протягом наступних двох днів. З'єднувальні роботи почалися негайно, і очікується, що вони будуть завершені наприкінці квітня. Цей відносно довгий проміжок часу пов'язаний з дощем, який робить зварювання особливо складним. Решта будівельних робіт на стартовому майданчику йде за графіком. Фундамент столу для запуску завершений на 60% і буде завершений до кінця липня. Роботи над зонами зберігання пального почалися нещодавно, і ситуація зараз наступна: 40 UDMH: завершено стіну захисного ставка зберігання завершено роботи почалися нещодавно. Також розпочалося будівництво резервуару для охолодження збереженого пального та кондиціювання приміщень у фундаменті столу для запуску та центрі керування. Дороги вже прокладені і будуть завершені наприкінці літа. Будівля збірочного цеху була відремонтована, а приймальний огляд повинен відбутися в травні.(4)

Improvement of launch-vehicle performance A number of improvements to the Ariane launcher presently under development enable future users to be offered an increase in payload from 1500 kg to at least 1600 kg in transfer orbit. It can be assumed that in this way Atlas-Centaur-type performance will be reached, allowing satellites like Intelsat-V to be launched by Ariane. This can be achieved without compromising the overall programme schedule and the cost-to-completion ceiling. The following measures have been taken or are under study: improvement of third-stage HM7 engine, resulting in higher specific impulse increase in propellant loading of second-stage L33 increase in propellant loading of first stage L 140 optimisation of the first-stage engine nozzle shape, reSUlting in higher specific impulse and increased thrust at lift-off lightening of second and third-stage H8 equipment.

The Apex Programme

Within the framework of the development of the European Ariane launch vehicle, managed by the French national space organisation CN ES, four development flight tests are foreseen in the years 1979 and 1980. The primary objective of these flights will be to qualify the launch vehicle. In addition, the Ariane Programme Board of ESA has authorised the flight of passenger experiments on board the second, third and fourth development flights (coded L02, L03 and L04). This offer of free flight opportunity has been transmitted to the Agency's Member States, and to organisations in non-Member States with whom ESA has close links. Each experimenter is, however, responsible for all passenger-related costs. In response to the offer, 93 experiments were proposed, most of which were non-autonomous in the sense that a spacecraft structure, attitude-control system, power supply, etc. still had to be provided. This approach of providing non-autonomous experiments turned out to be fairly costly for the experimenter and if maintained would have led to considerable underemployment of the capacity available on the development flights. A more acceptable approach of selecting autonomous passengers from already-assembled spacecraft remaining in hand from earlier space programmes has now been substituted.

ARIANE LAUNCH VEHICLE The recently improved nominal payload capability for Ariane is the injection of 1600 kg into a geostationary transfer orbit, with perigee 200 km, apogee 36 000 km, and inclination 10.50 • As the launch vehicle carries additional measurement equipment on these development flights, in each stage as well as in the payload area, the volume and mass budgets available for the passenger are slightly reduced and the orbital parameters are somewhat different, in particular the inclination which is 17.70 for L02 because of the longer visibility range required for extended launch-vehicle telemetry reception during qualification flights.

POSSIBLE PASSENGER MISSIONS As the number of proposed passenger experiments exceeds the number of development flight opportunities, a selection has to be made taking into account the interest that exists in a particular mission in the ESA Member States, as well as the feasibility of the mission itself and any funding constraints.

FLIGHT OPPORTUNITY L02

The L02 launch has been assigned essentially to scientific passenger pay loads, to be chosen from the following:

COS-B Second Flight Model with' an additional boost motor to raise the apogee provided by Ariane, the scientific experiments being the same as on the first COS-8 mission.

GEOS Second Flight Model, with the same scientific experiments as will be carried on the first flight model next year.

COSARI: For this mission, the existing COS-8 spacecraft would be modified to carry scientific experiments to investigate the magnetosphere and the plasma, as well as experiments on relativity (gravitational red shift). extreme ultraviolet emission and the measurement of the diffuse infrared sky background.

APPLE: This passenger, proposed by the Indian Space Research Organisation (ISROl. comprises a small experimental geostationary telecommunications satellite with a C-band communication transponder (4-6 GHz). An Effective Isotropic Radiated Power sufficient to cover the I ndian subcontinent is foreseen (at least 30.2 dBW).

AMSAT: The International Amateur Radio Satellite Organisation has proposed, through its German affiliate , an amateu r radio satellite of the so-called 'Oscar-type' , weighing 68 kg and carrying two redundant transponders receiving in the VHF and transmitting in the UHF band, allocated to the use of amateurs.

COS-B, COSARI and GEOS are considered the principal candidates, with APPLE and AMSAT as additional passengers that could be accommodated in a lateral position within the payload fairing . The Indian telecommunications satellite is also being considered as a possible additional passenger for the L03 flight .

FLIGHT OPPORTUNITY L03

The proposed passengers for this flight are:

OTS: An OTS platform procured on a marginal cost basis, additional to the platforms for the main OTS-Marots programme, could carry a modified telecommunications payload operating in the 4/6 and 2.5 GHz frequency bands.

Meteosat: This mission could provide (i) a complementary se ri es of experiments to the basic Meteosat system, by having the two satellites in orbit simultaneously (giving increased image rate and stereoscopy). or (ii) in-orbit redundancy to increase operational reliability, if necessary, after 1981, or (iii) experiments with new spacecraft- or observation-instrument -related technologies.

Symphonie: The Franco-German Symphonie programme has proposed the Symphonie protoflight model as an APEX passenger. This model, built with flight-proven hardware, was originally to be launched on the last pre-operational flight of the now superseded ELDO Europa-II launch vehicle. Its launch on L03 would allow the Symphonie programme to be extended beyond 1980/81, when the two flight models then in orbit will be approaching the end of their expected lifetime. The investment made in the ground segment could also then provide a return for an extended period.

Canadian Satellite derived from CTS: The Canadian authorities have informed ESA of their interest in taking advantage of APEX flight opportunity to extend the CTS programme, the basic assumption being that the CTS Engineering Model can be refurbished to serve a:: a telecommunications payload platform . One payload option might be a UH F transponder designed to test telecommunication systems required for the far north of Canada after 1980. None of the four satellites proposed as principal passengers for the L03 flight will occupy the full capacity of the launch vehicle, and flight of a combination of any two might have been a theoretical solution. However, in the light of the major modifications that would be needed structurally and in the launch procedures, the Agency does not intend to examine this possibility further. Nevertheless, accommodation of the Indian APPLE satellite as an additional passenger on the L03 flight is still being seriously investigated; as this satellite's configuration has not yet been frozen, its design could take into account the constraints of a combined launch.

FLIGHT OPPORTUNITY L04

The L04 vehicle, the last of the qualification launches, might be of help in preparing for the Agency's second generation of appl ications satell ites. Both ESA and its Member States are studying how to make the best use of this launch. Several satellite configurations offering high power and long lifetime are under investigation and could be flight tested and qual ified on L04. A technical payload, such as a direct television or an advanced telecommunications mission, could be accommodated, and discussions are still in progress between ESA and its Member States.

CONCLUSION

At the moment, the Agency is performing an accommodation study to investigate the feasibility of pairing two satellites - one of the proposed principal passengers and one of the proposed additional passengers - for the L02 and L03 launchers. Final selection of the L02 and L03 passengers should take place in Autumn 1976, with the results of the above studies and an assessment of the Agency's scientific and/or technical interests in the principal programmes serving as a basis for decision. A detailed mission analysis will subsequently be conducted prior to final project-like implementation of the passengers chosen. Use of Ariane development fl ights to launch low-cost autonomous flight hardware in the way described will be to the benefit of the overall space activities of all organisations participating in the APEX programme.

Покращення продуктивності ракет-носія. Численні покращення ракети-носія Аріан, які зараз розробляються, дозволяють майбутнім користувачам отримати можливість запускати вантаж від 1500 кг до щонайменше 1600 кг на трансферну орбіту. Можна припустити, що таким чином буде досягнуто продуктивність, схожу на Атлас-Сентавр, що дозволить запускати супутники, подібні до Інтелсат-V, за допомогою ракети Аріан. Цього можна досягти, не порушуючи загального графіку програми та обмежень витрат на завершення. Були прийняті або розглядаються наступні заходи: покращення двигуна третьої ступені HM7, що призводить до збільшення конкретного тягового опору; збільшення завантаження палива другої ступені L33; збільшення завантаження палива першої ступені L140; оптимізація форми сопла двигуна першої ступені, що призводить до збільшення конкретного тягового опору та збільшення тяги при злітанні; полегшення обладнання другої та третьої ступені H8.

Програма Apex

У рамках розробки європейської ракети-носія Аріан, яку керує французька національна космічна організація CNES, передбачено проведення чотирьох випробувальних польотів у 1979 та 1980 роках. Основною метою цих польотів буде кваліфікація ракети-носія. Крім того, Комітет програми Аріан ЄКА дозволив проведення польотів пасажирських експериментів на другому, третьому та четвертому випробувальних польотах (позначених як L02, L03 та L04). Ця пропозиція безкоштовної можливості польоту була передана членам Агенції та організаціям у країнах, які не є членами ЄКА, з якими ЄКА має тісні зв'язки. Кожен експериментатор несе відповідальність за всі витрати, пов'язані з пасажирами. На пропозицію було подано 93 експерименти, більшість з яких були незалежними в тому сенсі, що структура космічного апарату, система контролю орієнтації, джерело живлення і т. д. все ще мали бути надані. Виявилося, що цей підхід надання незалежних експериментів виявився досить витратним для експериментатора, і якби його продовжувати, це призвело б до значного недопрацювання доступних можливостей на випробувальних польотах. Тепер його замінили більш прийнятним підходом вибору автономних пасажирів серед вже зібраних космічних апаратів, які залишилися від попередніх космічних програм.

РАКЕТА-НОСІЙ "АРІАН" Недавно покращена номінальна можливість навантаження для "Аріана" - це введення 1600 кг на геостаціонарну трансферну орбіту, з перигеєм 200 км, апогеєм 36 000 км та нахилом 10,50 °. Оскільки ракета-носій несе додаткове вимірювальне обладнання на цих розвідувальних польотах, в кожному етапі, а також у зоні навантаження, обсяги та масові бюджети, доступні для пасажира, трохи зменшуються, і орбітальні параметри трохи відрізняються, зокрема нахил, який становить 17,70 для L02 через більший діапазон видимості, необхідний для прийому розширеної телеметрії ракети-носія під час кваліфікаційних польотів.

МОЖЛИВІ МІСІЇ ПАСАЖИРІВ Оскільки кількість запропонованих пасажирських експериментів перевищує кількість можливостей для розвідувальних польотів, потрібно зробити вибір, враховуючи інтерес, що існує до певної місії в країнах-членах ЄКА, а також можливість самої місії та будь-які обмеження фінансування.

МОЖЛИВІСТЬ ПОЛЬОТУ L02

Польот L02 в основному призначено для наукових пасажирських навантажень, які будуть обрані з наступних:

COS-B Друга модель польоту з додатковим підвищувальним двигуном для підняття апогею, наданим "Аріаном", наукові експерименти такі ж, як на першій місії COS-8.

GEOS Друга модель польоту з такими самими науковими експериментами, які будуть проведені на першій моделі польоту наступного року.

COSARI: Для цієї місії існуючий космічний апарат COS-8 буде модифікований для проведення наукових експериментів з вивчення магнітосфери та плазми, а також експериментів з відносністю (гравітаційний зсув), екстремальним ультрафіолетовим випромінюванням та вимірюванням розсіяного інфрачервоного небесного фону.

APPLE: Цей пасажир, запропонований Організацією індійських досліджень у космосі (ISRO), включає в себе невеликий експериментальний геостаціонарний телекомунікаційний супутник з комунікаційним транспондером С-діапазону (4-6 ГГц). Передбачається ефективна ізотропна випромінювальна потужність, достатня для покриття індійського півострова (принаймні 30,2 дБВт).

AMSAT: Міжнародна організація аматорських радіосупутників запропонувала через своє німецьке підрозділ, аматорський радіосупутник так званого типу 'Оскар', вагою 68 кг і з двома резервними транспондерами, які приймають у діапазоні VHF та передають у діапазоні UHF, виділені для використання аматорами.

COS-B, COSARI та GEOS вважаються основними кандидатами, а APPLE та AMSAT - додатковими пасажирами, які можуть бути розміщені у бічному положенні всередині обтічного кожуха. Індійський телекомунікаційний супутник також розглядається як можливий додатковий пасажир для польоту L03.

МОЖЛИВІСТЬ ПОЛЬОТУ L03

Запропоновані пасажири для цього польоту:

OTS: Платформа OTS, придбана на маржинальних витратах, додатково до платформ для основної програми OTS-Marots, може перевозити модифіковану телекомунікаційну навантаження, що працює в діапазонах частот 4/6 та 2,5 ГГц.

Meteosat: Ця місія може забезпечити (i) додаткову серію експериментів для базової системи Meteosat, маючи два супутники на орбіті одночасно (що забезпечує збільшену частоту зображень та стереоскопію), або (ii) резервування на орбіті для збільшення операційної надійності, якщо це буде необхідно після 1981 року, або (iii) експерименти з новими технологіями, пов'язаними з космічним апаратом або обладнанням для спостережень.

Symphonie: Франко-німецька програма Symphonie запропонувала протопольотну модель Symphonie як пасажира APEX. Ця модель, побудована з перевіреним у польоті обладнанням, спочатку мала бути запущена на останньому попередньо-експлуатаційному польоті застарілої вже ракети-носія ELDO Europa-II. Її запуск на L03 дозволив би програмі Symphonie продовжити діяльність після 1980/81 року, коли два польотні моделі, які тоді будуть на орбіті, наближатимуться до кінця їх очікуваного терміну служби. Інвестиції, зроблені в наземний сегмент, також можуть повернути прибуток протягом тривалого періоду.

Канадський супутник, похідний від CTS: Канадські органи повідомили ЄКА про свою зацікавленість скористатися можливістю польоту APEX для розширення програми CTS, базовою умовою є те, що інженерна модель CTS може бути відновлена для використання як телекомунікаційна платформа . Одним з варіантів навантаження може бути УКВ транспондер, призначений для тестування телекомунікаційних систем, необхідних для далекого півночі Канади після 1980 року. Жоден з чотирьох супутників, запропонованих як основні пасажири для польоту L03, не займе повний обсяг ракети-носія, і польот комбінації будь-яких двох може бути теоретичним рішенням. Однак, у зв'язку з великими модифікаціями, які були б потрібні структурно та в процедурах запуску, Агентство не має наміру досліджувати цю можливість далі. Тим не менш, можливість розміщення індійського супутника APPLE як додаткового пасажира на польоті L03 все ще серйозно вивчається; оскільки конфігурація цього супутника ще не була заморожена, його проектування може враховувати обмеження спільного запуску.

МОЖЛИВІСТЬ ПОЛЬОТУ L04

Верхній ступінь L04, останній з кваліфікаційних запусків, може бути корисним у підготовці до другого покоління супутників для застосувань Агентства. Як ЄКА, так і її країни-члени вивчають, як зробити найкраще використання цього запуску. Декілька конфігурацій супутників, що пропонують високу потужність та довгий термін служби, перебувають у стадії розслідувань і можуть бути протестовані в польоті та пройти кваліфікацію на L04. Технічне навантаження, таке як пряме телебачення або передова місія зв'язку, може бути розміщено, і обговорення між ЄКА та її країнами-членами все ще тривають.

ВИСНОВОК

На даний момент Агентство проводить дослідження розміщення для вивчення можливості поєднання двох супутників - одного з запропонованих основних пасажирів та одного з запропонованих додаткових пасажирів - для ракет-носіїв L02 та L03. Остаточний вибір пасажирів L02 та L03 має відбутися восени 1976 року, з результатами вищевказаних досліджень та оцінкою наукових та/або технічних інтересів Агентства у відповідних програмах як основою для прийняття рішення. Після цього буде проведено докладний аналіз місії перед остаточною реалізацією обраних пасажирів. Використання розвитку ракети "Аріан" для запуску недорогого автономного обладнання для польотів, як описано вище, буде корисним для загальних космічних діяльностей всіх організацій, що беруть участь у програмі APEX.(5)

First firing of the propulsion system The Ariane first-stage propulsion system comprising four Viking II engines underwent its first firing test on 17 November 1976 at SEP, Vernon . This firing test, developing a nominal thrust of 244 t, was terminated after 59 s - instead of 75 s as planned - because of an abnormal rise in the ambient temperature. The latter was found to be due to the failure of a small pipe associated with a pressure sensor, hence to the combustion of UDMH flow inside the bay. This test also showed an unexpected acoustic level, which is probably peculiar to the test stand.

Taking into account the multiple objectives associated with this test: qualification of Europe's largest test stand built by SEP at Vernon simultaneous start-up and functioning of the four Viking 11 en - gines commissioning of the water reservoir functioning of the hot-gas pressurisation system. the operation can be considered a complete success.

Data processing has yielded very satisfactory results. The hot-gas pressurisation system proved to operate nominally. The pressure surge phenomenon - which occurs at the end of propulsion during the sudden cut-off of propellant flow in the tubes - did not exceed the admissible values. The pressure peak at start-up, its level during the test and the temperature rise in the engines remained within the predicted limits.

Second-stage propulsion system The G1 and G2 firings of the second stage propulsion system - a Viking III engine with a truncated diverging section (the contoured diverging section Viking IV being designed for the flight model) - performed at DFVLR (Hardthausen) were also highly successful (130 s of functioning developing a thrust of over 60 t). Subsequent tests will make use of the flight-model cold-gas pressurisation system. Simultaneously, vacuum testing of the Viking IV engine will start at Hardthausen. Three firings of the complete H M 7 engine As significant as the combined firing of the first stage's engines was the functioning of the third stage's complete HM 7 engine during three tests on 30 November, 1 and 8 December (1710 s of operation with excellent results) . This was the cryogenic engine for which a major test facility had been built at Vernon. Development of the combustion chamber at MBB has been completed and qualification tests are due to start in January 1977. This chamber provides a specific impulse of 440 s (compared with the 434 s required by the specification) and modifications are being made in order to improve its performance. The next major phases in the development of the third stage will be the propulsion bay test with full tanks scheduled for January 1977 and the first vacuum testing of the complete HM 7 engine planned for 15 March. Fairing Dynamic testing of the fairing (diameter: 3.2 m, height: 8.65 m). development of which was assigned to Contraves, has just begun with two separation tests for the rear cone. Detailed results are not yet known, but a quick analysis of the first test revealed a satisfactory lateral velocity and successful separation. The fairing jettison tests will start in March in ESTEC's large vacuum test chamber.

Перше випробування системи приводу Перша ступінь ракети Аріан, що складається з чотирьох двигунів Вайкінг II, пройшла перше випробування 17 листопада 1976 року у SEP, Вернон. Це випробування, під час якого розвивалася номінальна тяга 244 т, було припинено після 59 с - замість запланованих 75 с - через аномальний підйом температури навколишнього середовища. Виявлено, що це сталося через відмову маленької трубки, пов'язаної з датчиком тиску, тому що відбувалася спалювання потоку УДМГ всередині бака. Це випробування також показало неочікуваний акустичний рівень, ймовірно, притаманний саме стенду для випробувань.

Беручи до уваги кілька цілей, пов'язаних з цим випробуванням: кваліфікація найбільшого в Європі стенду для випробувань, побудованого SEP у Верноні одночасний запуск і функціонування чотирьох двигунів Вайкінг 11 введення в експлуатацію резервуара для води функціонування системи підвищення тиску гарячого газу. Операцію можна вважати повністю успішною.

Обробка даних дала дуже задовільні результати. Система підвищення тиску гарячого газу виявилася працювати номінально. Феномен підвищення тиску, який виникає в кінці приводу під час раптового вимикання потоку пального в трубах, не перевищував допустимих значень. Піковий тиск під час запуску, його рівень під час випробування та підйом температури в двигунах залишалися в межах передбачених лімітів.

Система приводу другої ступені G1 і G2 запуски системи приводу другої ступені - двигун Viking III з обрізаним розширювальним відрізком (контурний розширювальний відрізок Viking IV розроблений для літакової моделі) - виконані в DFVLR (Хартхаузен) також були дуже успішними (130 секунд функціонування з розвитком тяги понад 60 т). Наступні випробування будуть використовувати систему підвищення тиску холодного газу для літакової моделі. Одночасно розпочнеться вакуумне випробування двигуна Viking IV в Хартхаузені. Три запуски повного двигуна H M 7 Ще більш значущим, ніж спільний запуск двигунів першої ступені, було функціонування повного двигуна третьої ступені HM 7 під час трьох випробувань 30 листопада, 1 та 8 грудня (1710 секунд роботи з відмінними результатами). Це був кріогенний двигун, для якого було побудовано великий випробувальний майданчик у Верноні. Розробка камери згоряння в MBB завершена, і кваліфікаційні випробування повинні розпочатися у січні 1977 року. Ця камера забезпечує конкретний імпульс 440 с (порівняно з 434 с, які вимагаються у специфікації), і вносяться зміни для покращення її продуктивності. Наступні важливі етапи в розвитку третьої ступені будуть випробування бака для приводу з повними баками, заплановані на січень 1977 року, і перше вакуумне випробування повного двигуна HM 7, заплановане на 15 березня. Обтічник Динамічні випробування обтічника (діаметр: 3,2 м, висота: 8,65 м), розробленого Contraves, щойно розпочалися з двома випробуваннями розділення для задньої конусної частини. Детальні результати ще не відомі, але швидкий аналіз першого випробування показав задовільну бічну швидкість та успішне розділення. Випробування відкидання обтічника розпочнуться у березні у великій вакуумній камері в ESTEC.(6)

Performance

Ariane's capacity for launch from Kourou into geostationary transfer orbit (200- 36000 km, inclination 10.5") has been upgraded to 1700 kg . The previous guaranteed performance was 1600 kg - itself an improvement on the initial capability (1500 kg) . This further improvement resulted from the latest assessment of the launcher characteristics which in - dicated an expected performance limit of 1763 kg . As a result of this assessment by CNES, based on reliable data (actual weights of structure and equipment and measured specific impulses of the engines) and making due allowance for contingencies, ESA is henceforth guaranteeing to users a minimum performance of 1700 kg in geostationary transfer orbit. The improvement in performance resulted from a reduction in the weight of the 3rd Stage structure, an increase in the 2nd Stage fuel loading (UDMH and N20 4 ) from 33 to 34 tonnes, an in - crease in the specific impulse of the HM7 engine to 435 s, and a better knowledge of the launch vehicle's aerodynamic properties after tests carried out by ONERA on scale models. Ariane will now be able to place in geostationary orbit satellites of 925-965 kg depending on the performance of the ABM used. Ariane's vibratory characteristics

The sinusoidal vibration test standards for payloads to be launched by the operational vehicle have been improved as a result of:

- a re -analysis of the POGO loop stability and of the potential efficiency of the POGO correcting system installed on the 1 st and 2nd Stages which led to a reduction of the excitation level between 10 and 35 Hz (along the longitudinal axis) from 4 g to 1.5 g.

- results from specific tests carried out on the 3rd Stage of the launcher's dynamic mockup which showed that the vibrational stresses to be applied to the payloads during sinusoidal vibration tests above 100 Hz could be suppressed along all three axes.

Interface Review

An Interface Review held from 20 January to 7 March with the participation of ESA and CNES representatives resulted in a critical ana - lysis of the current version of the User 's Manual and various recom - mendations for future issues, and it was decided to issue a new edition in July 1977. This new manual will incorporate most of the proposed changes. Some of the changes however require a longer preparation and will not be introduced until January 1978.

The Review also provided both users and launcher authorities with a better understanding of their respective constraints. Accordingly the next issue of the Manual should be a great improvement and a valuable tool for the promotion of Ariane.

Launch base facilities The first of the three new tracking stations for reception of the launcher's S-band telemetry data was ready for acceptance in mid - March at ChiHeauroux (see opposite) . These three telemetry receiving stations are to be installed, respectively, on Mount Galliot (near Kourou) in August 1977, on Mount Montabo (near Cayenne) in November 1977 and on the launch range at Barreira do I nferno, Natal (Brazil) in March 1978.

Each station is equipped with a 10-m antenna providing a 43 dB gain. Together with the Ascension Island facilities, they will ensure the re - ception of telemetry data throughout the launcher 's trajectory from the launch pad at Kourou to the point of orbital injection.

Natal Down-range Station Civil engineering work has been in progress since January on the Brazilian launch range at Barreira do Inferno (CLFB 1) for new facilities required both for Ariane use and the Brazilian programme.

This Natal station will be used for acquisition and distribution of all tracking and telemetry data transmitted by the launch vehicle.

This will involve the use of the existing Bearn radar for tracking purposes (see photograph, page 50) and the setting - up of an entirely new receiving station for the telemetry data. Виконання

Можливість Аріана запуску з Куру у геостаціонарну трансферну орбіту (200-36000 км, нахил 10,5") була підвищена до 1700 кг. Попередня гарантована продуктивність становила 1600 кг - що само по собі є покращенням початкових можливостей (1500 кг). Це подальше покращення було результатом останньої оцінки характеристик ракети, яка вказувала на очікуваний ліміт продуктивності в 1763 кг. В результаті цієї оцінки CNES, на підставі надійних даних (фактичні ваги конструкції та обладнання та виміряні конкретні тяги двигунів) та з урахуванням можливих непередбачених обставин, ESA відтепер гарантує користувачам мінімальну продуктивність 1700 кг у геостаціонарній трансферній орбіті. Покращення продуктивності стало результатом зменшення ваги конструкції 3-го етапу, збільшення завантаження палива 2-го етапу (UDMH та N20 4) з 33 до 34 тонн, збільшення конкретної тяги двигуна HM7 до 435 с, а також кращого розуміння аеродинамічних властивостей ракети після випробувань, проведених ONERA на макетах умовних розмірів. Аріана тепер зможе розмістити на геостаціонарній орбіті супутники вагою від 925 до 965 кг в залежності від продуктивності використаного АБМ. Вібраційні характеристики Аріани

Стандарти вібраційних тестів з синусоїдальними коливаннями для корисних навантажень, які мають бути запущені операційним пристроєм, були покращені завдяки:

• переаналізу петлі стабільності POGO та потенційної ефективності встановленої системи корекції POGO на 1-му та 2-му етапах, що призвело до зменшення рівня збудження між 10 та 35 Гц (уздовжної вісі) з 4 g до 1,5 g.
• результатів конкретних випробувань, проведених на динамічному макеті 3-го етапу ракети, які показали, що вібраційні напруження, які мають бути застосовані до корисних навантажень під час випробувань з синусоїдальними коливаннями вище 100 Гц, можуть бути прибрані уздовж усіх трьох осей.

Аналіз інтерфейсу

Аналіз інтерфейсу, який відбувся з 20 січня по 7 березня з участю представників ESA та CNES, призвів до критичного аналізу поточної версії Посібника користувача та різних рекомендацій для майбутніх питань, і було вирішено випустити нове видання у липні 1977 року. Цей новий посібник буде включати більшість запропонованих змін. Деякі зміни, однак, потребують більш тривалої підготовки і не будуть введені до січня 1978 року.

Аналіз також надав як користувачам, так і владі ракети кращого розуміння їх відповідних обмежень. Відповідно, наступне видання Посібника повинно бути великим покращенням та цінним інструментом для просування Аріани.

Функції стартової площадки Перша з трьох нових станцій відстеження для прийому телеметричних даних S-діапазону ракети була готова для прийняття наприкінці березня в Ші-Шору (див. протилежно). Ці три станції приймання телеметрії обладнані антеною діаметром 10 м з коефіцієнтом підсилення 43 дБ. Разом з установками на острові Асунсьйон вони забезпечать прийом телеметричних даних на протязі траєкторії ракети від стартової площадки в Куру до точки орбітального введення.

Станція в Наталі для прийому даних на відстані. Будівельні роботи тривають з січня на бразильській стартовій площадці в Баррейра-ду-Інферно (CLFB 1) для нових установок, які потрібні як для використання Аріани, так і для бразильської програми.

Ця станція в Наталі буде використовуватися для отримання та розподілу всіх даних відстеження та телеметрії, переданих ракетою.

Це включатиме використання існуючого радару Берн для відстеження (див. фотографію, стор. 50) та створення зовсім нової станції для прийому телеметричних даних(7)

Hardware manufacture under way for L 01 flight test In less than two years - on 15 June 1979 to be precise - the first flight of the Ariane launcher will take place. Manufacture of the hardware for this first L01 launcher began during the first quarter of 1977, and assembly of the three stages is scheduled for completion for the beginning of launcher integration in November 1978. The status of the programme, which is running on schedule, can be summed up as follows:

- all the structures have been developed and most of the qualification tests have been completed

- engine development has been completed and the qualification tests are in progress

- the long-duration tests for the engines have been completed

- the propulsion-bay system tests at stage level are under way

- the electrical mock-up tests have started

- the dynamic mock-up tests are now complete and analysis of the data obtained confirms the calculational models of the first modes.

- so far no technical problem has emerged which might lead to a slip in schedule.

In the area of propulsion, a major step has been achieved regarding the first stage, with the highly successful cluster firing tests of the propulsion system, totalling 402 s during 10 tests on 4 bays. For the second stage, the 'battleship' test series was continued very satisfactorily with the G M1 11 test lasting 137 s by integrating the flight configuration equipment.

For the second stage, a series of three consecutive firings took place, integrating the pressurisation system, the servo motors and the flightconfiguration roll -control system. The results were excellent. For the third stage, chamber qualification tests are complete and show a specific impulse of 444 s, i.e. 10 s better than specified. This is a remarkable result.

In addition, the propulsion -bay tests started in early May on stand PF42, reproducing the sequence of preparations for the stage at the Guiana Space Centre (CSG) . The synchronised sequence was validated . The first firing test was carried out on 21 July.

The first flight-configuration cryogenic stage has just been set up on stand PF43 and the filling tests have started.

The 15th vacuum ignition test of an HM7 engine on the PF41 stand was carried out, lasting the nominal 570 s.

First full-scale fairing separation test performed Ariane's fairings are designed to protect the vehicle's payloads against aerodynamic heating and other harmful environmental effects during ascent, within a useful payload volume of about 50 ma Their general layout is as follows: an aft cone with an interface diameter to the equipment bay of 2.6 m opens to 3.2 m maximum diameter and is followed by a cylindrical section 4 m high that terminates in a front cone section with a spherical nose. The overall height of the fairings is 8.6 m.

Ariane fairings provide radio transparency by several means. The aft cone consists of a glass fibre - Kevlar sandwich structure. I nail metallic parts, i.e. cylinder and front cone, access doors or cutouts with RFT-characteristics can be accommodated. Standard payload connectors on a support of variable length, to cope with a large variety of satellite size and connector accommodation, are available. A second device can be provided for dual-launch purposes.

Fairings can now be delivered with an acoustic blanket for those payloads susceptible to lift-off and aerodynamic noise. Fairing separation will take place at an altitude of some 110-140 km, determined by the payload requirements and the trajectory flown . A parallel separation mode similar in principle to the separation system used many times for Thor- Delta fairings will be employed.

To simulate mode and altitude of fairing separation, ground tests under vacuum are needed and therefore ESTEC's large Dynamic Test Chamber (DTC) was chosen for the Ariane tests.

After a series of separation tests at component and panel level to develop the vertical and horizontal separation systems, followed by a series of three separation tests using a fairing aft cone only, and after detailed dynamic analysis, development activities have now converged towards full-scale separation tests. The main objectives are ·to study separation behaviour, to validate test procedures, to learn how to handle a complex structured measurement system and to run through a process of complete data evaluation.

The first of these took place on 18 June in the DTC. The fairings separated successfully and all separation systems functioned correctly.

A number of lessons were learnt. Preparation and execution was found to be good. The clearance around the launcher was found to be better than predicted, although detailed analysis of trajectories and correlation with mathematical models is still going on.

The structure failed locally and modifications are already being made to avoid further failures, particularly rivet failures . The fairings are already undergoing a refurbishing process in preparation for a second test at the end of October. This test will be the first of a series of two qualification tests using the same fairing model (SM1).

After having executed the next separation test, static qualification will finally begin in November of this year, when the second qualification model SM2 will be available. Should results be as expected, the go-ahead will then be given for manufacture of the first L01 flight unit.

Виготовлення апаратури наразі триває для першого польоту L 01. Менше ніж за два роки - а саме 15 червня 1979 року - відбудеться перший польот ракети-носія Аріан. Виготовлення апаратури для цієї першої ракети L01 розпочалося в першому кварталі 1977 року, а збірка трьох ступенів запланована на завершення для початку інтеграції ракети в листопаді 1978 року. Статус програми, яка ведеться за графіком, можна узагальнити наступним чином:

• всі конструкції розроблені, і більшість випробувань на відповідність пройдені
• розробка двигунів завершена, а випробування на відповідність тривають
• випробування двигунів тривають
• випробування системи пропульсійного відсіку на рівні ступенів тривають
• випробування електричного макету розпочато
• випробування динамічного макету зараз завершені, і аналіз отриманих даних підтверджує розрахункові моделі перших режимів.
• до цього часу не виникло жодної технічної проблеми, яка могла б призвести до затримки в графіку.

У сфері пропульсії був досягнутий важливий крок щодо першого етапу, з високоуспішними тестами кластерного запалювання пропульсійної системи, загалом 402 с під час 10 тестів на 4 відсіках. Для другого етапу серія тестів "лінійки" продовжувалася дуже задовільно з тестом G M1 11 тривалістю 137 с шляхом інтеграції обладнання конфігурації польоту.

Для другого етапу відбулася серія з трьох послідовних запалювань, з інтеграцією системи підвищення тиску, серводвигунів та системи керування коченням у конфігурації польоту. Результати були відмінні. Для третього етапу випробування на відповідність камери завершені і показують специфічний тяговий коефіцієнт 444 с, тобто на 10 с краще, ніж вказано. Це вражаючий результат.

Крім того, випробування пропульсійного відсіку розпочалися на початку травня на стенді PF42, відтворюючи послідовність підготовки до ступеня на Космічному центрі Гвіани (CSG). Синхронізована послідовність була підтверджена. Перше випробування запалювання відбулося 21 липня.

Перший криогенний ступінь у польотній конфігурації тільки що було встановлено на стенді PF43, і розпочалися випробування на заповнення.

15-е вакуумне запалювання двигуна HM7 на стенді PF41 було проведено, триваючи номінальні 570 с.

Перше повномасштабне випробування розділення обшивки виконано. Обшивки Аріан розроблені для захисту навантажень транспортного засобу від аеродинамічного нагрівання та інших шкідливих навколишніх впливів під час підйому, у корисному об'ємі навантаження приблизно 50 м3. Їх загальна конструкція така: задня конусна частина з діаметром інтерфейсу до обладнаного відсіку 2,6 м відкривається до максимального діаметра 3,2 м і слідує циліндрична ділянка висотою 4 м, яка завершується передньою конусною частиною з кулястим носом. Загальна висота обшивок становить 8,6 м.

Обшивки Аріан забезпечують радіопрозорість кількома способами. Задня конусна частина складається з сендвіч-структури з скловолокна та кевлару. Металеві частини, тобто циліндр та передня конусна частина, двері доступу або вирізи з характеристиками RFT можуть бути розміщені. Стандартні роз'єми навантаження на підтримці змінної довжини, щоб впоратися з великим різноманіттям розмірів супутників та розміщення роз'ємів, доступні. Другий пристрій може бути наданий для подвійних запусків.

Обшивки тепер можна доставляти разом із звукоізоляційним ковдрою для тих навантажень, які вразливі до шуму під час зльоту та аеродинамічного шуму. Розділення обшивки відбудеться на висоті близько 110-140 км, визначеній вимогами до навантаження та траєкторією польоту. Буде використаний паралельний режим розділення, схожий за принципом на систему розділення, використану багато разів для обшивок Thor-Delta.

Для моделювання режиму та висоти розділення обшивки необхідні наземні випробування в умовах вакууму, тому для випробувань Ariane була обрана велика динамічна тестова камера ESTEC.

Після серії випробувань розділення на рівні компонентів та панелей для розробки вертикальних та горизонтальних систем розділення, наступила серія з трьох випробувань розділення, використовуючи лише задню конусну частину обшивки, а після детального динамічного аналізу роботи розробки зараз збігаються до випробувань розділення на повному масштабі. Основні цілі - вивчити поведінку розділення, підтвердити тестові процедури, вивчити, як працювати з складною структурованою системою вимірювань та пройти процес повного оцінювання даних.

Перше з цих випробувань відбулося 18 червня в DTC. Обшивки успішно розділилися, і всі системи розділення працювали правильно.

Було вивчено кілька уроків. Підготовка та виконання виявились добрими. Дозвіл навколо ракети виявився кращим, ніж передбачалося, хоча детальний аналіз траєкторій і кореляція з математичними моделями ще триває.

Конструкція місцево пошкодилася, і вже робляться модифікації для уникнення подальших вад, зокрема, вад з заклепками. Обшивки вже проходять процес відновлення перед другим випробуванням наприкінці жовтня. Це випробування стане першим із серії двох кваліфікаційних випробувань за тією ж моделлю обшивки (SM1).

Після проведення наступного випробування розділення статична кваліфікація нарешті розпочнеться у листопаді цього року, коли буде доступна друга кваліфікаційна модель SM2. Якщо результати будуть такими, як очікувалося, тоді буде дано дозвіл на виготовлення першої льотної одиниці L01. (8)

The series of ten tests on four firststage propulsion bays have just been completed, the final test lasting 30 s and making a total of 404.5 s in this configuration . The series began in November 1976, with the following main objectives:

- investigation of the behaviour of the propulsion -bay in the dynamic and thermal environment created by the simultaneous functioning of the four engines

- optimisation of the transients affecting the tank -pressurisation system and the start-up/ cut-off control system on start- up and shut-down of the four engines

- checking of the steady regime

- verification of the operation of the Pogo-correction systems and the nozzle actuators

- checking of the procedures for operating and qualifying the ground facilities prior to the stage tests. During these tests. flight-standard equipment was progressively integrated. with the one exception of the two main nitrogen tetroxide and U D M H tanks. which were of the 'battleship ' type in order to allow the pressurisation system to be adjusted. The capacity of the battleship tanks allowed the engines to be run for some 88 s compared with 145 s with the flight-standard ones.

A major problem apparent during the first test was rupturing of the propellant-circuit seals, as a result of the dynamic environment induced inside the bay. A number of corrective measures enabled the long-duration tests to be conducted to propellant depletion. These same bays were also fired several times in succession in order to optimise the adjustments.

It was satisfactorily demonstrated that all the objectives had been achieved, and it was thus possible to introduce the flight-standard tanks for the series of tests on the complete stage.

The first stage was transferred from the Launcher Integration Site at Les Mureaux to the Vernon Test Centre in a pressurised container 21 m long, 5 m wide and 6 m high. It constituted an out-of-gauge load, for which special arrangements were necessary.

It might be recalled that the first stage is to undergo four series of development tests between November 1977 and May 1978, followed by three series of qualification tests between September 1978 and March 1979.

Серія з десяти тестів на чотирьох двигунах першого етапу щойно завершилася, останній тест тривав 30 секунд і загалом у цій конфігурації тривав 404,5 секунди. Серія почалася в листопаді 1976 року з наступних основних цілей:

• дослідження поведінки двигуна у динамічному та термічному середовищі, створеному одночасною роботою чотирьох двигунів
• оптимізація перехідних процесів, що впливають на систему тискіння бака та систему керування запуском / вимиканням при запуску та вимиканні чотирьох двигунів
• перевірка сталого режиму
• перевірка роботи систем корекції Пого та приводів форсунок
• перевірка процедур експлуатації та підтвердження здатності наземних засобів перед етапними випробуваннями. Під час цих випробувань стандартне обладнання для польотів поступово інтегрувалося. за винятком двох основних баків з азотною тетроксидом та U D M H, які були типу "лінкор", щоб дозволити налаштувати систему тискіння. Ємність баків лінкор дозволяла працювати двигунам протягом близько 88 секунд у порівнянні з 145 секундами для стандартних для польоту.

Однією з основних проблем, виявлених під час першого тесту, було розривання ущільнень пропелерного контуру внаслідок динамічного середовища всередині бака. Декілька коригувальних заходів дозволили провести тестування тривалості до вичерпання палива. Ті самі баки було також кілька разів постріляно поспіль для оптимізації налаштувань.

Було успішно продемонстровано, що всі поставлені цілі були досягнуті, і таким чином стало можливим впровадження стандартних для польоту баків для серії тестів на повному етапі.

Перший етап було перевезено з майданчика інтеграції ракет на Ле Мюро до Центру випробувань у Верноні в тисненому контейнері завдовжки 21 метр, шириною 5 метрів і висотою 6 метрів. Це було негабаритне навантаження, для якого були необхідні спеціальні заходи.

Можна пригадати, що перший етап має пройти чотири серії випробувань розвитку між листопадом 1977 року та травнем 1978 року, за якими слідуватимуть три серії випробувань кваліфікації між вереснем 1978 року та березнем 1979 року.(9)

All three stages of the launcher have now reached the phase of completestage testing .

First stage The success of a series of propulsion - bay tests totalling 404.5s of operation made it possible to introduce flightconfiguration tanks and to begin the series of development tests on the stage. The first test, carried out at the Vernon test centre on 13 December, was satisfactory, in spite of premature cut-off. The test was stopped after 111 s, instead of the planned 150s, following erosion of the graphite throat of one of the four Viking engines. The functioning of the hot gas pressurisation system in relation to the flight tanks was shown to be nominal, as was the stage's dynamic behaviour.

Second stage The first test of the second-stage in flight configuration took place successfully at the DFVLR test centre at Hardthausen on 31 January. The stage, equipped with its pogocorrection system, functioned as planned for 138s, with swivelling of the engine.

The firing was terminated on UDMH depletion. All parameters were nominal. The test allowed the compatibility of the propulsion system with the flight tanks to be checked, as well as the performance of the pressurisation system.

Third stage Following the long-duration firing (lasting 470s) of the propulsion system in its 'battleship version', the first test of the third stage in flight configuration was carried out successfully at the Vernon test centre on 10 January. The stage functioned nominally for the exact planned duration of 250s. A second test took place on 2 February; it lasted 550s as planned, which corresponds pratically to the duration of third-stage powered flight (570s). The purpose of the tests was to check compatibility of the propulsion bay with the tanks in flight configuration, with particular reference to the pressurisation systems. The successful completion of these events shows that the technology has now been mastered.

Qualification of the stage-separation system The fourth test, which took place on 15 December, completed the qualification phase for the pyrotechnic separation systems. For both stage separations (1/2 and 2/3). the results have allowed the pyrotechnic design to be validated and correct behaviour of the equipment to be verified.

Delivery of the check out system to the Ariane launch site The launch-vehicle checkout system intended for the Ariane launch site at Kourou (French Guiana) was delivered in December, and installation began immediately. In the meantime, validation operations using static simulators are in progress. An identical checkout system has been installed since last November at the Launcher I ntegration Site at Les Mureaux, France, where the electrical mockup tests are currently under way.

Усі три етапи стартового пристрою вже перейшли до фази завершальних випробувань.

Перший етап Успішність серії випробувань віддушовування, загальна тривалість яких склала 404,5 секунди, дозволила ввести конфігурацію польоту для баків та розпочати серію випробувань розвитку на цьому етапі. Перший випробування, яке відбулося 13 грудня в центрі випробувань у Верноні, було задовільним, незважаючи на передчасне відключення. Випробування було зупинено після 111 секунд, замість запланованих 150 секунд, через ерозію графітового горла одного з чотирьох двигунів Viking. Робота системи гарячого газу для тиснення відносно баків польоту показала себе як нормальна, так само як і динамічна поведінка етапу.

Другий етап Перше випробування другого етапу в конфігурації польоту відбулося успішно в центрі випробувань DFVLR у Хардтхаузені 31 січня. Етап, обладнаний системою погокорекції, працював згідно з планом протягом 138 секунд з обертанням двигуна.

Вогонь було припинено при вичерпанні УДМГ. Усі параметри були нормальними. Випробування дозволило перевірити сумісність системи тяги з баками польоту, а також роботу системи тиснення.

Третій етап Після тривалого вогню (470 секунд) системи тяги в "броненосцівській версії", перше випробування третього етапу в конфігурації польоту пройшло успішно в центрі випробувань у Верноні 10 січня. Етап працював нормально протягом точно запланованої тривалості 250 секунд. Друге випробування відбулося 2 лютого; воно тривало 550 секунд, як заплановано, що практично відповідає тривалості польоту з третього етапу (570 секунд). Метою випробувань було перевірити сумісність багажного відділення з баками в конфігурації польоту, зокрема систем тиснення. Успішне завершення цих подій показує, що технологію тепер володіють.

Кваліфікація системи розділення етапів Четверте випробування, яке відбулося 15 грудня, завершило фазу кваліфікації піротехнічних систем розділення. Для обох розділень етапів (1/2 та 2/3) результати дозволили підтвердити піротехнічний дизайн та перевірити правильну роботу обладнання.

Постачання системи перевірки на стартовому майданчику Аріана Система перевірки ракети, призначена для стартового майданчику Аріана в Куро (Французька Гвіана), була доставлена у грудні, і встановлення почалося негайно. Тим часом, проводяться операції валідації з використанням статичних симуляторів. Така ж система перевірки була встановлена з листопада минулого року на майданчику інтеграції стартового пристрою в Ле-Мюро, Франція, де в даний момент тривають електричні тестування макету. (10)

15

The Ariane programme, which consists of a development phase followed by an operational (utilisation) phase, is one of the most important programmes ever undertaken by the European Space Agency. It originated from a proposal by the French Government in December 1972, and a subsequent decision in July 1973 by ESRO's Council, that the development phase would be carried out as an Agency programme. To this end, legal arrangements were concluded between ESRO and the participants in the programme (all the Member States of the Agency, except for Ireland and the United Kingdom, the latter participating in the programme under a bilateral agreement with France), and also between ESRO and CNES, which was to manage the programme for ESA. The aim of the Ariane programme is to provide Europe with an independent launch capability as from late 1980 . onwards. Surveys show that the World market for satellites in the period 1981 - 1990 could be substantial and that Europe is in a position to acquire a significant number of launches, particularly in the applicationssatellite field. It would already seem that European missions alone will guarantee two launches per year, a figure that could be raised to four or five by vigorous sales promotion at European level. After five years of development work and less than a year before the first test flight, scheduled for June 1979, the programme is proceeding according to plan . The period between now and this first flight will be devoted mainly to 'finishing touches' and qualification of the launcher stages and equipment bay, as well as the commissioning of the Ariane launch base, at Kourou in French Guiana. The other three test flights are scheduled for December 1979, May 1980 and October 1980. They will complete the qualification of Ariane, which will then be available for operational launches from late 1980 onwards. The decision by the ESA Council to begin manufacture of a series of five launchers, to be available from late 1980, constitutes another important milestone. The size of this first series is justified by the number of approved programmes making use of Ariane launches: namely, the scientific satellite Exosat (launch foreseen for second quarter 1981), the Marecs-B satellite (second quarter 1981), the ECS-1 satellite (fourth quarter 1981), all of which are ESA programmes, and the French earth -observation satellite SPOT, scheduled for launch in late 1983. Other European missions, such as the second unit of the ECS programme and the heavy satellite, H-Sat, will very probably also rely on Ariane launchers. In addition, ESA has undertaken a campaign aimed at markets outside Europe. For example, after very thorough studies conducted in conjunction with the I NTELSAT Organisation, ESA has tendered for the launch by Ariane of units 5, 6 and 7 of Intelsat-V, competing against the NASA tender for launch services using the Space Shuttle. These tenders are now being evaluated, and a decision is due to be made this autumn . The importance of this decision, not only for Ariane but for European aerospace industry as a whole, is self-evident. If demand from the users justifies, steps will be taken in conjunction with industry to either extend manufacture of the first series or undertake a second one. By way of conclusion, I should like to take this opportunity to thank all those who have contributed, and I am sure will continue to do so, to the success of the programme - the Delegations to the Ariane Launcher Programme Board and in particular its Chairman, Mr van Eesbeek, the many European industrialists, the staff of CN ES, which manages the programme, and the ESA staff, all of whom have worked unstintingly on this programme.

The general architecture of the Ariane vehicle reflects a concern to make maximum use of existing skills and hardware and of technologies that have either been qualified previously in other programmes or represent a minimum development risk. From the various possible alternatives, a three-stage launcher based on conventional technology was chosen which, though heavier at lift-off, would also be less costly to develop and to produce than one involving major technical innovations. Optimisation of the propellant masses of each stage and the general dimensions of the vehicle led to different diameters being adopted for the first stage and for the two upper stages. However, adoption of a bulb- shaped fairing provides sufficient usable volume to house large satellites. The only advanced propulsion technology employed is that of the third stage, where use of liquid hydrogen and liquid oxygen has allowed the requisite performance to be obtained without unduly increasing vehicle dimensions.

DESCRIPTION OF THE LAUNCHER

Ariane is a three -stage launcher with a total height of 47.4 m and weighing 208 t at lift-off, 90% of the mass being constituted by propellant. The structures and the payload account for about 9% and 1 % of the total weight, respectively.

FIRST STAGE (L 140) The first stag e weighs 13.2 t empty and has a height of 18.4 m and a diameter of 3.8 m. It is equipped with four Viking -V engines, developing a thrust of 2445 kN (about 245 t) at lift-off and 2745 kN in vacuo (specific impulse 281 .3 s) . The 147.5 t of propellants (UDMH and N2 0 4 ) - of which 815 kg remain unburnt after 145 s of flight - are contained in two identical tanks of 15 CDV - 6 steel connected by a cylindrical skirt.

The whole of the lower part of the stage, which comprises the four engines, the thrust frame, the water tank, propulsion -system accessories, the cowlings and the fins for aerodynamic stabilisation during atmospheric flight, constitutes the propulsion bay of the L 140 stage. The four turbopump engines and 54 bar combustion chambers are mounted symmetrically on the thrust frame and can be swivelled in pairs about two orthogonal axes to provide three-axis control. The propellant feed is provided by a turbopump with a flow of 270 kg/s at a pressure of 70 bar. Propellant intake is effected through a radial injector. The refractory steel chamber has a single wall cooled by propellant film injected along the wall and is fitted with a bell-shaped nozzle with a graphite throat. To avoid pump cavitation, the tanks are pressurised to about 5 bar by the gases produced by the generator associated with each engine. This generator uses the same propellants as the main engine, but the gases are cooled by water injection. They power the turbine of the turbopump unit of each Viking engine as well as the hotgas motor of each of the four actuators that command the swivelling of the engines.

SECOND STAGE (L33)

The second stage weighs 3.22 t empty (without the interstage and the jettisonable acceleration rockets) and has a height of 11 .6 m and a diameter of 2.6 m. It is equipped with a Viking -IV engine which develops 717 kN of thrust in vacuo (specific impulse 294 s) . The engine is attached to the tapered thrust frame by a gimbal with two degrees of freedom for pitch and yaw control, roll control being effected by auxiliary jets fed with hot gas tapped from the stage gas generator. The two propellant tanks are· of A -Z5G aluminium alloy, have a common bulkhead, and are pressurised with gaseous helium (3.5 bar) ; they contain 34.2 t of propellants (U DM Hand N2 0 4 ) of which 137 kg remain unburnt after 138 s of flight.

THIRD STAGE (H8)

The third stage is the first cryogenic stage developed in Europe. It weighs 1.157 t empty and is 9.08 m high and 2.6 m in diameter. It is equipped with an H M -7 engine, which develops a thrust of 60 kN (specific impulse 441 s). The two tanks, which contain 8.23 t of propellants (liquid hydrogen and liquid oxygen), of which 67 kg are unburnt after 570 s of flight, are made of A-Z5G aluminium alloy, chosen for its good behaviour at the temperature of liquid hydrogen ( - 20 K), and have a common bulkhead (two walls separated by a vacuum) . They are clad with an external thermal protective layer of KIE§gecell to prevent the heating of propellants. The hydrogen and oxygen tanks are pressurised in flight by gaseous hydrogen and helium, respectively. The engine turbine, which is fed by the gases from a generator, drives the oxygen pump at 12000 rpm and the hydrogen pump at 60 000 rpm. The combustion chamber is of the regenerative-cycle type. The walls are cooled by the circulation of fuel through a network of slots adjacent to the chamber before its admission to the axial injector, which consists of 90 elements arranged in concentric circles. The construction of the body of the combustion chamber relies on original technology (developed by MBB) the patent of which is also applied in the United States in constructing the combustion chamber of the Space Shuttle's main engine: the cooling slots are milled in a copper casting and then covered with an electrolytic deposit of nickel. The chamber opens into a bell-shaped nozzle consisting of spiralled tubes of Inconel cooled by the circulation of hydrogen which vaporises in the slots. The engine is attached to a tapered thrust frame, gimbalmounted for pitch and yaw control, roll control being provided by auxiliary nozzles which eject gaseous hydrogen. Stage separation is achieved by linear shaped pyrotechnic charges located in the aft skirts of the second and third stages. The stages are moved apart by retro-rockets mounted on the lower stage and acceleration rockets attached to the upper stage.

EQUIPMENT BAY

The equipment bay weighs 319 kg and is 2.66 m in diameter and 1 .15 m high. It is mounted on the third stage and houses the vehicle's electronic equipment, supports the payload and provides the attachment points for the fairing . In it are centralised all the launcher functions (sequencing, guidance, flight control, tracking, destruction and telemetry) . Only the actuating and executive systems are distributed among the stages. The guidance and control system, based on a digital computer and an inertial platform, detects the vehicle 's attitude and measures its acceleration, From this information and the instructions contained in the guidance program, the computer provides navigation and guidance. It generates and transmits attitude-correction commands to the stages of the launcher. An analogue flight-control unit mixes the required attitude deviations (provided by the computer) with the information supplied by the rate gyros, After filtering the structural and liquidsloshing critical modes, the autopilot sends commands to the hydraulic actuators that swivel the engines, as well as commands for the opening of the roll-control jets of the second stage and the attitude-control jets of the third stage. At the end of flight, when the velocity corresponding to the desired orbit is attained, the computer orders propulsion cut-off. The precision thus obtained is of the order of 5 m/s for a velocity of more than 10000 m/so

FAIRING

The fairing weighs 826 kg and is 3.2 m in diameter and 8.65 m high (external dimensions) . It protects the payload during the ascent through the atmosphere and is jettisoned during the flight of the second stage, at an altitude of about 110 km. It consists of two half-shells of aluminium with a boat-tail section of laminated material which is radio transparent. Its useful volume of 35 m3 allows large geostationary satellites of the Intelsat-V or H -Sat type to be carried, or two medium size satellites mounted one above the other in the 'Ariane dual launch system ' (Sylda) .

PROGRESS IN THE PROGRAMME

In rather less than a year (in June 1979) the first Ariane flight test will take place at the Guiana Space Centre. Now that the ground testing of the various launcher subsystems is coming to an end and the stage and system tests are in full swing, it is perhaps appropriate to review the programme as a whole.

SYSTEM STUDIES

The general studies for the launcher (trajectory, performance, aerodynamics, flight control, stresses, dynamics, guidance, flight mechanics, thermal and acoustic) have been conducted in several iterations as hardware definition and tests have progressed. They are now largely completed, and work in this area is concentrated on adjustments and preparations for the first flig ht (L01 ) and its exploitation . These studies, carried out by Aerospatiale (programme system integrator), checked by CNES and, in critical cases, confirmed by independent studies carried out by ON ERA, have been the subject of intense critical examination and verification . Ground testing and flight simulation have allowed the coherence of subsystem specifications and the various margins and dispersions to be checked . It is noteworthy that the general performance currently guaranteed to Ariane users is a satellite mass of 1700 kg in transfer orbit (200/ 36000 km). appreciably more than the 1500 kg in the original specification. This remarkable improvement achieved during the development, a rare occurrence in the aerospace field , results mainly from the care with which the hardware specifications were drawn up.

SYSTEM TESTS

The purpose of the system tests is to check the results or to confirm the assumptions of the systems studies, and also to check the coherence between the various subsystems. They comprise six series of tests: aerodynamic tests launcher dynamic tests launcher electrical tests guidance and flight-control simulation tests launcher/launch-site compatibility tests stage-separation tests. The aerodynamic tests have already been completed. They have enabled the launcher's flight dynamics to be verified and the loads and the thermal fluxes to which the structures will be subjected to be checked. The launcher dynamic tests have also been completed . They have allowed the frequencies and characteristics of the various longitudinal and transverse vibration modes to be verified, with a view to predicting and correcting the launcher's vibratory behaviour in flight, These tests were wide ranging and were conducted on real and complete structures representing the launcher configuration at various characteristic moments of flight,

The aims of the launcher electrical tests were firstly to verify the electrical compatibility of the launcher's various items of equipment and of its ground checkout facilities, and secondly to develop the automatic checkout software used for integration in France and for launchings in G uiana. They were conducted using a real equipment bay, a checkout system corresponding to that to be used at the time of launch, and electrical stage simulators. They were started more than a year ago and were completed in July. They took longer than was initially foreseen because of the lengthy validation of the very many automatic checkout modules required during launch. The guidance and flight-control simulation tests have allowed the in-flight behaviour of the guidance and flight-control system for various dimensioning trajectories to be simulated, and dispersions to be studied. They are conducted with real elements of the system up to and including the hydraulic actuators for the nozzles and the nozzles themselves and a computer unit enabling hybrid simulation (digital and analogue) of the launcher's behaviour during flight. These tests are also well advanced and will be completed this summer. The launcher/launch-site compatibility tests are in - tended, as their name implies, to verify the complex interfaces between the launcher and the launch site, to develop launch procedures and to train the appropriate teams. They comprise numerous simulation phases and exercises, the most important of which is erection of the launcher and the rehearsal of launch operations using a vehicle specially provided for the purpose. This operation is due to take place at the Guiana Space Centre in the period August to November. The propellant mock-up, consisting of the structures and almost all the propulsion equipment in flight configuration, left Aerospatiale's integration site at Les Mureaux in the second half of June. Lastly, the stage-separation tests, which have now been completed, have permitted the proper operation of the pyrotechnic devices that cut the structures joining the stages and ignite the retro -rockets of the lower stages and the acceleration rockets of the upper stages, to be verified.

THE STAGES

The first static firing of the first stage (L 140) took place on the SEP test stands at Vernon on 13 December 1977, and was followed by a second test on 9 March. Previously, a number of engine tests and ten propulsionbay tests had been carried out between 17 November 1976 and 29 September 1977. The only major technical problem still unresolved concerns the behaviour of the graphite throat of the combustion chamber of the Viking engines with which the first and second stages are equipped . Although this component satisfactorily withstood all the tests until the first test of the first stage, it proved insufficiently resistant during the long firing of the four nozzles together. This technical obstacle, which should be fairly easy to overcome by adopting other materials already qualified under the more stringent conditions prevailing in dry-propellant engines, nevertheless poses a tricky scheduling problem because of the time needed to qualify this major modification and assemble the flightstandard engines. The third development test in June will be followed by three qualification tests late in 1978 and during the first half of 1979. The configuration for the L01 launcher should be validated at the end of 1978. The first firing of the second stage (L 33) took place on the DFVLR test stands at Hardthausen (Germany) on 31 January. Two very satisfactory tests of nominal duration have now taken place. Qualification of the engine 's functioning in vacuo has been practically achieved. As for the first stage, an overall development test has still to be carried out. followed by three overall qualification tests. The first firing of the third stage (H8) took place at Vernon on 10 January and lasted 256 s. It was followed by three tests, including two of long duration (more than 500 s). Although this stage, which runs on liquid hydrogen and liquid oxygen, is by far the most complex, no major technical problem has so far arisen. Nevertheless, in view of the number of items of equipment in the stage, many more tests are still required in order to ensure the necessary reliability . In addition, the overall structure has successfully undergone qualification tests.

EQUIPMENT BAY

The equipment bay houses the electronic equipment needed for navigation, guidance, flight control, telemetry, generation of sequencer commands, and safety. A prototype of the bay was used for developing this assembly, and it was then handed over by MATRA to SN IAS with a view to overall testing of the electrical system. Currently, all the constituent electronic equipment has satisfactorily undergone its qualification testing, except for the inertial platform which is being qualified now.

FAIRING DEVELOPMENT

The fairing, which protects the payload during the vehicle's ascent through the atmosphere, has been developed by the Swiss firm of Contraves. Studies and preliminary tests have resulted in modifications to the system for separating the two half-shells. The whole assembly has been qualified following three separation tests in a large vacuum chamber.

ADAPTATION OF THE GUIANA SPACE CENTRE AND DOWN - RANGE STATIONS

Work on the Ariane Launch Site ('Ensemble de Lancement Ariane ' - E LA), which makes very extensive use of the original Europa-II facilities, is coming to an end and acceptance tests are being carried out. the first launcher being due at the ELA in early August. The work on adapting the CSG as a whole (logistics, safety and telemetry) will be completed this year, as will the links to the down -range tracking, telemetry and radar stations at Natal in Brazil and on Ascension Island (NASA and 000 stations) . It is planned to hold operational exercises for training the teams in the last quarter of 1978 and the first half of 1979.

CONCLUSION

Up to this very advanced stage in the ground testing programme, only routine technical problems have been encountered, with the one critical exception of the behaviour of the graphite throats of the Viking -engine combustion chambers. Because of its late appearance in one of the longest cycles in the development of the launcher, this problem has considerably upset the timetable for the first-stage qualification tests, which have consequently been delayed by nearly three months. Nevertheless, appropriate steps have been taken to ensure that the first flight test will be able to take place in June 1979 as planned .

Ariane Launch Performance

The original objective of the Ariane programme was the development of a launcher capable of placing a mass of 1500 kg into a transfer orbit leading to geosynchronous orbit. In 1976, following the definition phase, it was decided to increase the performance margin by some 100 kg over the guaranteed figure. To this end, the nozzles of the first-stage Viking engines were modified and, by reducing the ullage spaces, the masses of propellants carried in the first and second stages were increased by 5t and 1 t, respectively. In March 1977, when the results of the engine tests were known and the stages had been weighed, it proved possible to increase the guaranteed performance to 1700 kg, the difference between this and the calculated performance still! being some 100 kg (Fig.1 ).

MAIN MISSION

The Ariane launcher has been designed to place into transfer orbit (200/35850 km) a composite comprising a satellite and its apogee motor, the latter being used to place the satellite into its final geosynchronous orbit. The low latitude ofthe Guiana launch base (5.23°N) enables a particularly simple launch procedure to be used: each stage is ignited immediately after the burnout of the previous one, without an intermediate coast phase, thanks in particular to the long burn time of the third stage (some 10min). The launch azimuth selected allows the injection point for final orbit, namely the perigee point of the transfer orbit. to be located over the equator. The apogee of this orbit will therefore also be situated in the equatorial plane, which is a precondition for the circularisation manoeuvre using the apogee motor. The inclination of the transfer orbit thus attained is 9.65° (Fig. 2). By comparison, the procedure for launching American vehicles from Cape Kennedy is relatively more complex, because the location of the base (latitude 28.5°N) entails a coasting phase and a re -light of the vehicle's last stage, the inclination of the transfer orbit also being 28.5°. A mass of 1700 kg in transfer orbit, including a MAGE-III apogee motor (specific impulse 295 s), corresponds to a final mass in geosynchronous orbit of 1004 kg, or a satellite mass of 965 kg.

The parametric dispersions at injection into transfer orbit are as follows: Inclination Perigee altitude Apogee altitude ± 0.019° ± 0.45 km ± 43.2 km

SECONDARY MISSIONS

The general performances of the launcher for launch azimuths between 0° and 90° are shown in Figures 3 and 4. It will be seen that Ariane can satellise 4850 kg into a low orbit (perigee 200km, inclination 5.23°). For masses greater than 2500 kg, it will be necessary to revise the dimensioning of the upper part of the launcher. Heliosynchronous mission This type of orbit, the essential feature of which is that the local time of the subsatellite point remains constant, is used mainly for earth-observation missions such as meteorology and remote sensing . Such orbits are retrograde, with inclinations lying between 90° and 180°. As an example, for an 840 km circular orbit with an inclination of 98.76°, Ariane can satellise a mass of 2500kg.

Interplanetary missions (Fig. 5) The table below gives the performance of the launcher in various interplanetary-probe roles (declination of asymptote: - 5°). Mission Moon Venus Mars Mercury Payload mass (kg) 1000 790 660 0 Note: Departures from the recommended launch window may lead to a degradation in performance (some 20 kg for 20 min ). ADDITION OF A FOURTH STAGE

Exosat mission As already mentioned, the Ariane launcher has been designed to carry out a prime mission that does not call for third -stage re-light. For certain special missions requiring a long coasting phase before injection into final orbit, it will therefore be necessary to add a fourth stage to the launcher. The PO.7 stage, already used as the third stage of the Diamant B.P4 launcher, has been adopted as the Ariane fourth stage for the Exosat mission. For this mission, the perigee of the highly eccentric orbit (500/ 200000 km, inclination 75°) would be in the neighbourhood of the South Pole. The composite PO.7 + Exosat would therefore first be placed by the three-stage launcher into an elliptical transfer orbit of 75° inclination. After a ballistic phase of some 9000 s, PO.7 would be ignited and would then inject Exosat into the desired orbit. It will be seen that the use of PO.7 on Ariane appreciably increases the latter's performance in interplanetary mis - sions (Fig. 5). Out-of-ecliptic mission By using more complex procedures, e.g . swinging-by a planet such as Mars, Venus or the Earth, one can considerably increase the launcher's performance, if a much longer transfer time is accepted. A study has shown it would be possible to carry out an out-of-ecliptic mission by swinging by the Earth and then Jupiter and using a re-lightable fourth stage similar to the Symphonie apogee motor. The final satellite mass would be 350 kg for a heliocentric inclination of 68°, but the penalty would be a transfer time of some 600 days.

Increased Performance for Ariane

Provided development work and testing continue as successfully as they have to date and flight trials are successful. Ariane will be flightqualified by the end of 1980 and an operational European launch capability comparable to that of an Atlas-Centaur will then be available. In addition to the four vehicles for development flights, five Ariane vehicles have been authorised for production. Since the qualification of any major change in a vehicle's configuration takes several years. plans must already be made now for any further evolution in the Ariane basic design. THE NEED FOR A PERFORMANCE INCREASE In the past, many nonmilitary satellites have been designed for launch on Delta or Atlas-Centaur launch vehicles. The Delta-3914 launcher can transport about 900 kg and the Atlas-Centaur about 1850 kg into geostationary transfer orbit (from which the satellite can be injected into the geostationary orbit by an apogee boost motor). These two weight classes have become so much a standard for satellite designers that, despite the planned phase-out of the Delta and Atlas vehicles by the time the Space Shuttle comes into operation, a considerable number of such satellites will stiill have to be launched in the early 1 980s. NASA has made special provisions fortheir launch aboard the Space Shuttle (STS-Space Transportation System) . For Delta -class payloads, an upper stage called a Payload Assist Module (PAM) is being developed which can be used in conjunction either with the conventional launch vehicle or with the Shuttle. The Shuttle can launch several satellites in the course of one flight. Ariane is to be launched from the Guiana Space Centre in Kourou (French Guiana), which is some 5° north of the equator, into geostationary transfer orbits that have lower inclinations than those of US launchings from Eastern Test Range in Florida (approx. 28° north of the equator) . The velocity increment needed to rotate the Ariane transfer orbit into the equatorial plane and circularise it is therefore approximately 300 m/s less, which means in turn that less propulsion is needed from the satellite's apogee boost motor for a given satellite mass. The standard Ariane vehicle will be able to Inject a payload of at least 1700 kg into geostationary transfer orbit with a perigee of 200 km and an inclination of 9.5°. This performance is sufficient for the launching of an Atlas-Centaur class satellite (910 kg in geostationary orbit). Dual launchings would significantly improve the competitivity of Ariane for Delta-class payloads. The technical feasibility of a double launch of two independent satellites has already been established. but the present Ariane payload capability allows only the launch of an STS/PAM-D and a Delta-2914 class satellite, whereas after 1980 mainly satellites of the Delta 3910/PAM-D and STS/PAM -D class will be built. Two payload-capability levels have been identified that would allow Ariane to launch several convenient com - binations of Delta-class satellites in the 1980s, namely 1950 kg and 2300 kg .

PROGRAMME OBJECTIVES

Mission models show that the ability to launch 1950 kg would already be advantageous at the end of 1981 , whereas the need for the 2300 kg level is likely to arise in 1983. In view of this, work has to start in 1979 to allow vehicle modifications to be qualified and for the necessary modifications to be introduced into the production vehicles at the right time.

The choice of an appropriate complementary programme has to be based on the following criteria:

- high efficiency in terms of performance increase versus additional development cost

- short lead time

- compatibility with on -going development and facil - ities utilisation, and with production.

Consequently, the following short-term goals are being set for the Ariane performance increase:

- first possible launch of 1950 kg in early 1982

- first possible launch of 2300 kg in early 1983.

A POSSIBLE APPROACH

Studies have shown that the following steps constitute the most efficient means of achieving these objectives:

Step 1: Increase in payload capability to 1950kg, by

• an increase in Viking-IV and -V chamber pressures from 54 to 58 bar

• an increase in third -stage tank capacity from 8 to 10 t (tanks carrying only 9.5t in this step) .

Step 2: Increase in payload capability to 2300 kg, by

• the addition of two 6 t strap-on boosters on the first stage

• full third-stage tanks (10t)

A lengthening of the fairing by 300-500 mm to accommodate two STSjPAM-D satellites may also be en - visaged. Detailed studies are being carried out this year by CN ES and the industrial firms concerned, with a view to submitting a proposal in the autumn of 1978 to the ESA Council.

FUTURE DEVELOPMENTS

The plans and studies that have been discussed are concerned only with development work in the immediate future, up to the end of 1983, and are intended as an improvement package for the basic Ariane configuration already under development. The ESA Executive will soon start to consider the longer-term evolution of Ariane and the needs and possibilities for new developments in the launcher field. The most difficult problem will be to estimate the long-term trend in payload -capability re - quirements.

Sylda - A Dual-Launch System for Ariane

In view of the planned existence of several Deltaclass payloads - requiring approximately half of Ariane's available performance - in the early 1980s, and in order to increase Ariane's competitiveness, it has been decided to build a carrier structure capable of supporting and releasing two independent spacecraft. The development of such a dual-launch system, caJled 'Sylda' (Systeme de lancement Double Ariane), was initiated by ESA in June 1978. The Sylda system consists essentially of a support structure, three separation systems (one for each satellite and one for the structure itself). I n its baseline configuration (shown in Figure 1 ), it offers the upper satellite the same volume as is available to an STS/PAM-D* satellite; the lower satellite can be similar in size to Marecs. The mechanical and electrical interfaces at the separation plane are the same for both satellites, and identical to those specified for the PA M - D. Cut-outs or doors in the Sylda structure provide both access to the lower satellite and radio-frequency transparency for communications with this satellite. Separation is provided by classical Marman clamp bands released by pyrotechnically operated bolt cutters. The structure is made of an aluminium honeycomb core, covered with carbon-fibre layers. Use of this advanced technology results in a very light structure, the complete Sylda system weighing only about 165 kg. It will be bolted onto the conical structure of Ariane's equipment bay in place of the vehicle's standard payload adapter. As this standard adapter weighs 44 kg, the payload penalty resulting from the double-launch system is some 120 kg, leaving 1580 kg of mass in transfer orbit for the two satellites. The Sylda structure is dimensioned on the basis of rigidity rather than load requirements. In fact. longitudinal vehicle resonance oscillations are expected during flight mainly in two frequency bands (11 -18 Hz and 28-35 Hz) and the structure has to be so designed that the first longitudinal resonance frequency of the Sylda/satellite composite will lie between these bands. The first lateral resonance frequency of this composite is required to be about 6 Hz. The vibrational environment to which satellites carried on Sylda will be subjected is being studied using mathematical models of the launch vehicle, of Sylda and of two typical satellites; the first results are expected to be available in October 1978. The environment for the lower satellite is expected to be very similar to that predicted for single Ariane payloads.Sylda will be equipped with accelerometers to detect the in -flight level of vibrations near the satellite separation planes up to a frequency of at least 50 Hz. The measurements will be transmitted to ground via the launch vehicle 's telemetry system.

IN-FLIGHT OPERATION

After cut-off of Ariane's third-stage motor, the attitudecontrol system of this stage will orient the two Syldacarried satellites in any desired direction and spin the composite payload up to 5 to 10 rpm. The attitude-control system will then be switched off. The upper satellite will be the first to be separated, then the upper half of the Sylda structure, and finally the lower satellite, all separations taking place along the longitudinal axis of the third stage with relative velocities imparted by springs. Due to third-stage orientation errors, third-stage and satellite static and dynamic imbalances, and perturbations introduced by the separation mechanism itself, the orientation in space of the separated satellites deviates from the desired one and the satellite's spin axis moves on a nutation cone around the satellite's angular momentum vector, the apex of the cone being located at the satellite 's centre of gravity. For satellites with an inertia ratio of unity, the pointing error of the angular momentum vector is not expected to exceed 5 -r and the nutation cone half angle should not be larger than 2-4°. The specifications require that during separation of the upper part of the Sylda structure and of the lower satellite, the separating bodies must at no time come closer to each other than 30 mm. Analysis has shown that a spin rate of 10 rpm would reduce satellite pointing error and nutation, whilst a lower spin rate would help in respecting the clearance requirement. Furthermore, to avoid collision, the distance between all separated bodies must steadily increase. All in all, this means that both the initial spin rate and the relative separation velocities (and hence the spring forces). must be carefully chosen and optimised .

FUTURE DEVElOPMENT

It is planned to expand the cylindrical part of the Sylda structure by about 500 mm at a later date in order to provide the volume needed to house two STS/ PAM- D type satellites. The fairing volume would have to be expanded accordingly and Ariane's performance uprated to 2300 kg in transfer orbit. As readers of the article preceding this one will be aware, a complementary development programme along these lines is already under study.

The Ariane Launch Base

The Ariane Launch Base has been designed to allow both developmental and operational launches to be conducted from the Guiana Space Centre (,Centre Spatial Guyanais' or CSG) with the necessary safety and flexibility and with a frequency of up to four launches per year. During the initial development phase, there will be an average of two launches per year, but these may take place within three months of one another. The base's geographical position, close to the equator (5.2368°N), is a favourable factor when launching any satellite, and particularly so for a geostationary spacecraft. Launches can be made from the base on azimuths from - 10.5° through north to + 93.5°.

The Ariane Launch Base (Fig. 1) lies about 18 km from the town of Kourou, and consists essentially of two elements:

- the Ariane Launch Site ('Ensemble de Lancement Ariane ' - ELA), located within the Centre's perimeter and containing the specific facilities needed for final assembly, checkout and launch operations

- the Additional Ariane Facilities ('Moyens Complementaires Ariane' - MCA), comprising the Centre facilities or adaptations thereof and the down - range stations needed for carrying out the particular mission in question.

In addition, the base includes payload facilities, which will be made available to users for final preparation of their satellites. The ELA belongs to the European Space Agency, as do the other facilities and equipment financed and provided under the Ariane Programme. The agreement between ESA and the French Government concluded in May 1976, guarantees the Agency and its Member States freedom of access to and use of the ELA for the purposes of their programmes.

THE LAUNCH SITE

The ELA, which has been established by modifying the facilities of the forlT)er Europa-II launch site wherever possible, falls into fdur distinct zones: - a launch zone, ~iCh includes the launch -pad area and launch cent an assembly zone a propel/ant-suppo t zone, and a liquid-oxygen / liquid-nitrogen plant. The main areas of activity as far as preparation, checkout and operation of Ariane are concerned will be the launch - pad area, the servicing tower, inside which the launcher will be erected and connected to the ground equipment, and the launch centre, where nearly all checkout activities are to be conducted (Fig . 2) . The propellant-support zone and the liquid-oxygen/ liquid-nitrogen plant can be considered ancillary launch -site facilities. The launch-pad area is made up of the following main elements: The foundation, which supports the launch table on which the launcher rests (Fig. 3) . The launch platform, which provides the main access to the servicing tower, and protects the premises containing the equipment needed for launcher checkout. The forward erection area, in which the launcher elements are erected with the help of the servicing tower's travelling crane. The rear access ramp, which provides access to the launch platform, table and servicing tower. The servicing tower (Fig . 4), which is fully airconditioned and can enclose the launcher on its table as well as the umbilical mast. The umbilical mast, which acts as a support for the various connectors and umbilical arms, as well as for the cable ducts, pipework and umbilical junction boxes. The peripheral equipment and buildings, comprising essentially: • two plants for discharging, storing and transferring the toxic UDMH and N2 0 4 propellants • new toxic -propellant storage facilities, compris - ing two 115 m3 tanks for UDMH and a further two for N2 0 4 two facilities for storing and transferring cryo - genic propellants • a facility for storing 300 m3 of liquid nitrogen (five tanks) and producing gaseous nitrogen at high pressure • a plant for producing and storing the iced water needed for air-conditioning • a support-services building • a safety building, including a guardroom, and a full -time firefighting service for the whole site. The working area around the launcher is served by seven fixed platforms at convenient levels and a mobile platform, allowing any part of the payload to be reached . A room inside the servicing tower, between the mobile platform and the travelling crane, constitutes the clean area in which payloads will be prepared for mechanical assembly with the launcher. All working levels are served by a 1000 kg lift. All levels are also served by an internal and an external staircase, and there is a system for emergency evacuation. The launch centre, sited 200 m from the foundation, is a heavily protected, blockhouse-type building providing adequate protection for personnel during the final preparation, propellant filling and launch operations. It houses the checkout and control equipment for monitoring launcher-preparation operations.

The assembly zone consists essentially of: an assembly building, to which the Ariane stages are taken on arrival in Guiana, for visual inspection and preparation prior to erection in the tower a stores building for site and launcher spares an office building. The propellant-support zone, comprising two main storage facilities (Fig. 5). two garages for tanker trailers and a propellant-analysis laboratory, is used only for logistic support. The liquid-oxygen/liquid-nitrogen plant, located in the Kourou industrial zone, will produce these two cryogenic products to meet the requirements of the base.

THE ADDITIONAL FACILITIES

The additional Ariane facilities (MCA) embrace all the support needed for co-ordinating the preparation and execution of launch operations, including measurements that relate to the conduct of the launch and on-board experiments and ensure the safety of both personnel and property during operations. They consist essentially of the CSG measurement facilities (tracking, telemetry, etc.), located along the coastal strip near Kourou, with further facilities on the lies du Salut, and the associated downrange stations.

TRACKING

The CSG tracking system has four radars (two acquisition radars located 5 km from the ELA, and two high-precision tracking radars, one of which is located near Kourou and the other near Cayenne), two infrared tracking kine - theodolites, computers, data -transmission systems and display equipment. All these facilities are interconnected and they will allow Ariane's trajectory to be monitored by radar until 200 s after orbital injection and provide accurate optical attitude-restitution data from the moment of lift-off until the vehicle passes out of visual range. To allow the whole of the launch trajectory to be monitored, down-range stations in the form of the Natal radar facilities (Fig.6) located within the CLFBI (Brazilian launch base), and the Ascension Island radar facilities (000), from which the orbital injection of payloads can be observed, are also employed (Fig. 8).

TELEMETRY

The configuration of the CSG ground -telemetry stations and the down-range stations is such that it is possible: to acquire without interruption all the telemetry data transmitted in the E-band (2200-2290 MHz) by the launcher and the Ariane technological capsule ('Capsule technologique Ariane ' - CAT) during the launch phase, from the final preparatory operations on the pad, until 200 s after third-stage cut-off for vehicle-equipment-bay telemetry, and until loss of visibility by the station covering injection for the capsule telemetry to acquire all the telemetry information transmitted in the A-band (136-138 M Hz) by the technological capsule, from final preparations on the launch pad until fairing jettison to register all telemetry transmissions received on magnetic tape to reconstitute in real and deferred time the telemetry information received. The telemetry facilities include the Montagne des Peres (Kourou) and Cayenne-Montabo receiving stations, both operating in the 2200-2290 M Hz band, and the DianeIris station, which forms part of the satellite control network, operating in the 136-138 M Hz band. The Ariane programme also provides for setting up two new telemetry reception stations in Brazil; one, a mobile station, in the neighbourhood of Belem, and another in Natal at the CLFBI (Fig. 7) . The NASA telemetry reception station on Ascension Island will be used to cover the launcher's last powered phases and the placing of the satellite in orbit.

THE PAYLOAD FACILITIES

The payload facilities ('Moyens Charges Utiles' - MCU) are comprised of all the facilities made available by ESA to Ariane users for preparing their satellites, from arrival in Guiana until launch. They include the payload installations ('Installations Charges Utiles' - ICU) - buildings and equipment, specifically intended for preparing Ariane pay loads - and the various services provided for payload installation and satellite transport and preparation. To avoid any misunderstanding, it should perhaps be mentioned that the term 'payload' is used here to denote the mass that the Ariane launcher places in orbit for the customer. Generally speaking, this is the satellite, i.e. a mission module and a service module, and the propulsion module, which is an apogee motor in the case of a geostationary mission, but may also be a perigee motor (for certain missions there will be no propulsion module) . The ICU include several buildings dispersed geographically according to the activities to which they are assigned: Satellite-preparation and checkout building (B1) Propulsion-module plreparation building (B2) The satellite/propulsion-module integration building (B3) Propulsion-module storage building (B4) Upper platform of the servicing tower (platform 8) Storage buildings (B5) Launch centre (payload, technical and operational consoles) Operations centre (payload operational console) . Building B1 and the operations centre are located in the CSG technical centre, 5 km west of Kourou. Building B4 is located in the storage zone for solid -propellant thrusters, 8 km west of the technical centre. The other buildings, B2 and B3, are in the ELA assembly zone, 1.5 km south of the launch zone. The storage buildings are situated at various points within the Gu iana Centre. The ICU as a whole has been designed to handle satellites compatible with Ariane 's maximum performance. The distribution of the operations between several buildings allows the length of the launch campaign to be kept to a minimum even in the case of a double satellite launch in the same payload . Building B 1, intended for satellite preparation and checkout, provides a 420 m2 clean room which can be divided up for the simultaneous preparation of two smaller satellites for double launches. Buildings B4 and B2 provide for solid-propellant thrusters to be stored, refrigerated, submitted to X-ray inspection ar,d prepared before transport to building B3. The engines for both large and small satellites can also be processed in buildings B4 and B2. Provisi ~) n is made in building B3 for filling satellites w ith propellant and integrating the payload. In the case of a double launch, the two payloads will be integrated separately and then mounted on the Sylda* to form the payload composite, which is then placed in a special container. Communications between the satellite and its checkout system are provided by landline and radio relay. In principle, the checkout system is located in building B 1, but space is available in the launch centre for a checkout system of reduced dimensions along with the payload operation consoles. Telephone, interphone and television circuits can be set up on request between the various locations where payload activities are taking place, and all the buildings include suitable ancillary facilities such as offices, stores, laboratories and cloakrooms (specially equipped for propellant-work and clean -room clothing) .

Altitude-Simulation Testing of Ariane's Second-Stage Engine

When developing the propulsio n syst em f o r a particular stage of a launch veh icle, o ne of the major unknowns that has to be investigat ed is how that engine will function (ignition, performance, etc.) under the conditions prevailing at stageseparation altitude. Development testing must therefore be carried out by simulating the ambient conditions that the propulsion system will incur at its operating altitude. In the case of Ariane's second stage, this testing has been conducted at DFVLR's facilities at Hardthausen. The second stage takes over propulsion of the Ariane vehicle 52 km into the launch and carries it to a final altitude of 135 km in 138 s. During this period, the vehicle's velocity is increased from 2260 to 5160 m/ s, with a final acceleration of 45 m/ s2. After first-stage separation, the vehicle weighs 5 x 104 kg, and this falls to 1.6 x 104 kg as further propellants are used up. The same propellants are used as in the first stage, namely N20 4 (nitrogen tetroxide) and UDMH (unsymmetrical dimethylhydrazine). and they are contained in a tank with two compartments each of 15.5 m3 separated by a common bu Ikhead. Feed pipes carry the propellants to the Viking-IV turbopump engine, which delivers the thrust. The engine is started up via a control system which opens three main valves and activates the tank pressurisation systems. An eventual fall-off in acceleration signals propellant depletion and initiates third-stage separation.

THE VIKING-IV ENGINE

As one of the Viking family of rocket engines developed in France by SEP, the Viking -IV is specially designed to propel the upper stages of launch vehicles. In the case of Ariane, it propels the second stage, by producing a specified thrust level it provides the gases needed to power the flightcontrol servomotors and the roll-control system it provides hydraulic correction of Pogo effect (oscillations resulting from vibrational coupling between the launcher's structure and propulsion system). The main elements of the engine are: A turbopump, which augments the pressure of the propellants drawn from the stage tanks and injects them into the combustion chamber. A gimbal with two degrees of freedom, which allows the thrust vector to be swivelled through 4°, as required by the flight-control system. A thrust-unit ensuring combustion of the propellants and ejection of the gas produced. A gas generator, in which part of the main propellants are burnt and then cooled by injecting water, and which regulates the engine thrust. This gas is the energy source for the turbine, the servomotors and the roll-control system. The main valves, which control engine start-up and cut-off by admitting propellants and water. The Pogo correction systems (N 20 4 and UDMH). which absorb fluctuations in flow to the pumps to prevent their being coupled with the vibrations of the structure and the propulsion system. The operating principle and primary characteristics of the engine are shown schematically in Figure 1.

GENERAL TEST OBJECTIVES

The engine is being developed in three phases and the objectives of the accompanying test series can be summarised as follows: Turbopump tests, aimed at determining performances, and checking the mechanical endurance and settings of the gas generator. Complete engine tests under normal atmospheric ground conditions, with a nozzle whose exit pressure is suited to operation at atmospheric pressure. The main objectives of these tests are to verify the mechanical resistance of the hardware, the stability of combustion when running steadily, and the correct adjustment of the propellant mixture. The engine is used in this configuration for complete-stage tests, the purpose of which is also to study compatibility between the various subsystems. Testing of the turbopump and engine alone is carried out on the PF2 test stand at SEP, Vernon (France) and that of the complete stage at DFVLR. Hardthausen (Germany) . Engine tests with altitude simulation, the objectives being the measurement of propulsion performances, mechanical and thermal behaviour of the nozzle, and engine start-up under simulated flight conditions. This latter objective covers the stability of com - bustion, conditioned by the vaporisation of the propellants, the pressure at ignition and the build -up of chamber pressure produced by pressure in the tanks and the acceleration of the stage. The test facilities are required to provide an ambient pressure of 1.5 mbar at ignition, and during run -up a level allowing supersonic expansion . These tests are also carried out at DFVLR.

THE TEST FACILITIES

The test stand (Fig. 2) is an adaptation ofthe P4.2 stand at DFVLR, previously used for altitude-simulation testing of Europa-II 's third -stage engine. The vacuum chamber (1 ). in which the engine is installed, is 2.8 m square and 4.9 m high. The propellant tanks (2) , each with a capacity of 23 m3 , are installed above the vacuum chamber and arranged in a similar way to those of the second stage. They are pressurised by nitrogen so as to reproduce the pump-inlet pressure variations that occur in flight. The gases exit from the nozzle at a speed of 3000 m/ s, a stagnation temperature of 2900 K, and a static pressure of 230 mbar. They pass through a supersonic diffuser (3). which recompresses them to atmospheric pressure. To allow a vacuum to be created in the chamber before engine start-up, the diffuser outlet is blanked off by a membrane (4) which burns in the jet after ignition. The engine nozzle is surrounded by a screen (5) , cooled by a water jacket. intended to absorb the thermal energy radiated by the nozzle wall, the temperature of which reaches 11 OO·C. At the exit of the diffuser, the vertical gas jet is deflected obliquely (6) . The cooling system of the deflector and diffuser requires a coolant flow of 600 I/ s.

An 800 mm pipe (7) connects the diffuser to the steam ejector (8 and 9). which draws off the engine gases during the start-up phase. Part of the steam is condensed in an injection cooler (10) and the condensate falls into an underground sump. The flow of turbine exhaust gases is drawn off throughout the running of the engine by two steam ejectors (11 , 12). Part of the steam is also condensed by an injection cooler. The steam required for all nine ejectors is produced by two vaporisers (13 and 14). each of which comprises a rocketengine combustion chamber producing hot gases (by burning kerosene and nitric acid) into which water is injected to obtain steam.

CONDUCT OF THE TEST

When the mechanical and measurement facilities of the test stand have been duly prepared, a primary vacuum of a few millibars is first created in the chamber by a mechanical pump. The tanks are automatically filled with propellant and the sealing of all the propellant and vacuum circuits is checked . Two minutes before engine start-up, the steam generators are started and proper operation of the ejectors and generators is checked by observing the pressure, which must remain less than 1.5 mbar. When this level is reached, the valves between the ejectors and the vacuum chamber are opened. The automatic test sequence begins 40 s before engine start-up. Ignition is achieved by opening the engine 's main valves. During this phase, the automatic sequencer transmits commands to the test-stand systems: propellant circuits, diffuser cooling film, and measurement devices. At 0.2 s after the command to open the valves, a first increase in chamber pressure is noted; thereafter, in - creases of pressure to 25 mbar at the nozzle exit and 200 mbar in the combustion chamber denote the vaporisation of the N2 0 4 . The cooling due to this vaporisation produces a slight fall in pressure. Ignition results in an increase in chamber pressure from 0.2 bar to 3.5 bar, at 0.7 s. At that moment, the nozzle-exit pressure reaches 0.9 bar and the blanking membrane opens. The engine continues to run up and steadies out at 3 s; simultaneously, the suction due to the engine jet brings the vacuum-chamber pressure to 40 mbar when the nozzleexit pressure reaches 230 mbar. Depending on the test objectives, the chamber pressure and consequently the thrust or the mixture ratio, can be varied . After a maximum period of 180 s, cut-off is initiated by closing the main engine valves. During the test. 220 measurements are recorded and 70 commands transmitted. The functioning of the motor and the test stand are monitored by ten sensors which automatically terminate the test if certain parameters exceed prescribed limits. Figure 3 shows the Viking-IV engine being installed in the vacuum chamber on the test stand, linked by its gimbal to the thrust-measurement gauge and with the main valves connected to the propellant feed lines.

MAIN RESULTS

In the course of the Ariane development programme, eight tests of the type just described are due to be conducted, including two qualification tests. So far, five of these tests have been completed, three lasting more than 140 s, two short ones lasting less than 15 s (total 515 s) and one re-light. The three remaining tests should bring the total time in vacuum to 1000 s. All the main test objectives that were set for the altitudesimulation tests on the second stage have been successfully achieved, including measurement of an engine impulse of 295.8 s, nozzle-wall temperatures identical to those foreseen, and near-nominal combustion-engine pressures on start-up.

ARIANE

Ariane Launch Site The construction of the Ariane Launch Site (' Ensemble de Lancement Ariane' - ELA) which started in mid-1975 with the civil engineering work, has continued through 1976 and 1977, during which time nearly all the equipment has been supplied and installed. The early months of 1978 saw the installation of the last items of hardware (cryogenic arm and vehicle release system).

The vehicle checkout system, which arrived in Kourou in December 1977, was immediately installed in the Launch Centre and the checkout of interfaces was initiated. I ndividual acceptance tests of the ELA subsystems (phase-1 tests) started early in 1978; they will be followed by checkout system tests on dynamic simulators, and subsequently, from May to July 1978, with the ground facilities (phase-2 tests). The completion of the phase-1 tests in late June 1978 will mark the taking over by the site team of all the ELA facilities. In the first months of 1978, the various propellants and fluids needed for the vehicle were despatched to Guiana, including the liquid hydrogen, the sea transport of which was a World 'first'. The three stages and the various vehicle elements that constitute the propellant mock-up will leave Europe by ship in late June and arrive in Guiana in mid-July. The unloading of the stages and their despatch by road to the ELA will allow the launchvehicle transport and handling procedures in Guiana to be validated. The preparatory activity in Guiana will culminate with the vehicle erection tests and the propellant mock-up tests intended to qualify the procedures and the ground and onboard equipment needed for fill operations (phase-3 tests). These operations are scheduled for the period August to November. The Base validation tests will be concluded by a launch rehearsal, scheduled for the end of 1978 and involving all the elements of the ELA, the CSG facilities, and the down-range stations. Thus, by the beginning of 1979, the ELA will be ready for the first development launch (L01) of the Ariane vehicle. Additional Ariane facilities The preparation and completion of the Ariane additional facilities (Moyens Complementaires Ariane - MCA), which relate to the necessary adaptation of the CSG and the down-range stations, are proceeding normally. The electronics of the CSG radars have been renewed and the new E-band telemetry stations have now been set up at Montagne des Peres (near Kourou) and Montabo (near Cayenne). The last tests at these stations are nearing completion, The installation of a telemetry reception station with a configuration similar to that of the CSG stations is proceeding actively at Natal (Brazil) within the Brazilian launch base CLFB I. All the necessary steps have been taken to set up a mobile telemetry station in the Belem region (Salinopolis). Technical arguments relating to the development launches have in fact recently shown the usefulness of having telemetry reception facilities in an intermediate zone between the Cayenne-Montabo and Natal stations, Despite the short time available, the Salinopolis station will be ready early in 1979 to participate with the other telemetry stations in the operational qualification of the Launch Base system. The NASA and 000 tracking and telemetry stations on Ascension Island are ready to participate with the other Base facilities in carrying out global testing of the tracking and telemetry network, which will take place from late October until late January 1979. Operational qualification of the complete network will take place from February to midApril 1979, including the communication and data transmission facilities. The main instrument for this campaign will be the American GEOS III satellite; it will be tracked for this purpose when visible successively from Kourou, Salinopolis, Natal and Ascension.

Integration Testing of the Propulsion System of Ariane's First Stage

SEP, which is responsible for developing and testing the Ariane launcher's Drakkar propulsion system, has carried out separate development tests on the subsystems forming the propulsion bay. The second phase of testing, begun on 17 November 1976 with the first propulsion-bay firing, was designed to study the functioning of the integrated subsystems. Called 'cluster', or for brevity 'G', tests ('essais de groupement'), their purpose was to highlight mechanical, thermal or vibratory problems that might result from interaction between the various systems in configurations and under conditions similar to those to be expected during Ariane's flight.

TEST PROCEDURES

The propulsion bay, held in place by the same attachment points that will be used to mount it in the launcher, was supplied with propellants from two thick-walled tanks containing N2 0 4 and UDMH, respectively, and pressurised by hot gas provided by the bay 's own pressurisation system. The pipework was similar to that adopted for flight. This propulsion system only allowed firings lasting a maximum of 87 s because of the limited tank capacity, as against 147 s in flight. but the shorter duration was nevertheless adequate for obtaining thermal equilibria. The hardware was integrated progressively during the tests, the heavy thrust frame (5 t) used for the first bay being replaced by a light. flight-standard one (1 .6 t) for the second bay. The servomotors for swivelling the engines for flight-control purposes were introduced as from the third bay, which was once more mounted on a heavy thrust frame. The last bay, tested with a light thrust frame and Pogocorrection system, was of flight standard. The other subsystems - water tank, hot-gas pressurisation device, tank-bottom connections for propellants and pneumatic services, filling , draining and overflow valves for the tanks, command units, Viking-I! engines with conical nozzles (as opposed to the bell-shaped Viking-V nozzles to be used in flight) -were mounted on each bay, together with the hardware encasing the bay and the cowlings and heat shield. Originally, it was intended to use five bays for the cluster or G tests, each one being fired twice. The aim was firstly to study the re-light that might be necessary in the event of an aborted launch, and secondly to gain a better span of test conditions (ullage spaces and pressurisations in the tanks at the start of firing, burnout on propellant depletion, cut-off when closing the valves, etc.). After a first test without re-light, the tests on bays 2, 3 and 4 were finally conducted three times, and although only four bays were tested because of slippages in schedule, all test objectives were achieved .

TEST FACILITIES

A special test stand (PF 20) was built between 1974 and 1976 for the Drakkar propulsion -system, tests, twenty thousand tons of concrete being needed for its construction . In each case, the propulsion system under test is secured on a concrete' slab 3 m thick with a 5.5 mm square aperture through which the engine jets pass (Fig. 1). The latter are deflected 26 m below by the jet deflector, a buried flue with 800 t of uncooled protective lining. The 100 t securing device on the slab allows the flatness, parallelism/perpendicularity and spacing of the retaining jaws to be maintained to within 0.1 mm, to avoid setting up stresses in the thrust frame. The two removable 36500 I tanks, weighing a total of 52 t. are supported above the bay on four legs. The bay and tanks are protected by a metallic structure fitted with a travelling crane, which can handle the tanks, the bay, or the complete propulsion system with flight-standard tanks. This brings the total height of the structure above ground to 52 m. The associated propellant systems comprise two 125 m3 tanks for nitrogen tetroxide and U 0 M H, respectively, together with: the pumping systems needed for filling the tanks; nitrogen storage facilities with appropriate means for supplying at various pressures; safety facilities for injecting water, foam (intended to limit the evaporation of propellants in the event of an accidental leak) and products for neutralising propellant vapour or possibly dealing with polluted water; and facilities for decontaminating the propulsion system, which enable it to be moved and dismantled in complete safety. The measurement and control facilities are located together in an underground command post capable of withstanding an overpressure of 2 bar in the event of an explosion. In the course of a test, 500 parameters are processed there by a Mitra-15 computer, and a second such computer is used for conducting the test itself. Automatic, majority-logic monitoring of the main parameters enables the firing to be stopped if anything goes wrong. Fifty thousand digital measurements are made per second, and 52 parameters are recorded by means of a wideband tape recorder.

MAIN RESULTS

The first results revealed considerable noise at the bottom of the tank, between the four engines. Studies by SEP (modelling) , and tests by ONERA on the mock -up, revealed resonance of the aft cavity of the bay, between the heat shield and the cowl ings, when excited by the engine jets. They also showed that the phenomena would not recur in flight with the bell-shaped Viking -V nozzles. Between tests G3 -2 and G3-3 (test sequence and firing durations shown in Table 1), the Viking -II nozzles mounted on the bay were shortened by pyrotechnic cutting, and measurements during the latter test confirmed the influence of nozzle length on the noise, as predicted by ON ERA mock-up tests. The considerable noise and the simultaneous operation of the four engines led to a high vibration level which caused U D M H leaks during the first two firings. The propellant caught fire and the tests had to be stopped . Improved methods of locking the joints have obviated a recurrence of these incidents and two 87 s firings have since been carried out successfully. The G4 bay functioned for a total of 170.4 s without any problem. Leaks at the sliding joints between the engines and the U D M H valves occurred because of the vibratory environ - ment, and as from the third bay these joints were replaced by flexible connections, which have been proved satisfactory in use. An oil leak in this same third bay due to a pipework rupture in one servomotor (a weak point already earmarked for modification) and the failure of a speed regulator in another were also due to the vibratory environment. 'After modification, the servomotors behaved entirely satisfac - torily during the G4 bay's 170 s of running . During each test, the high acoustic level between the engines caused the flexible connections between the engines and the heat shield to deteriorate. This problem was only resolved during the first propulsion -system firing with the flight-standard tank in December 1977. Some of the cowlings also suffered from the acoustic environment, but a new model was successfully tested during the G4 test.

The first test had also revealed considerable oscillation of the chamber pressure on light-up, and this problem was resolved as from the G2-3 firing by staggering the times at which the tanks were initially pressurised . During the last G4 test series (Fig . 2). the hot-gas system for pressurising the tanks functioned unstably (without affecting the conduct of the firing), owing to a modification of the adjustments and the small ullage spaces in the tanks simulating the launch conditions in Guiana. Remedies that had been partially tried out as from the second G4 firing were successfully applied to the first propulsion-system firing with fIiIght-standard tanks. The G3-2 test, which ended on propellant (N 2 0 4 ) depletion, revealed a corrosion phenomenon in the tankbottom collector (flight-standard hardware) after the firing. This was found to be due to pollution of the propellants by water in the pressurisation gases and the tank bottom has since been modified to solve this problem which, of course, does not arise in flight. The G2 firing revealed a relatively large amount of unburned propellant (150 kg) on the N2 0 4 side. An antiresidual device was fitted as from test G3 and, by reducing the residual propellant mass to 30 kg, this has led to a 5 kg improvement in Ariane 's performance in transfer orbit. Failures of the valves of the bay's filling and draining system on the N2 0 4 side occurred one week after a first firing because of corrosion. This phenomenon does not occur after a short firing and is therefore not typical of the conditions in which the flight stages will operate, even in the case of an aborted firing . Generally speaking, the resistance of the hardware to propellants exceeded expectations (Fig . 3): the G4-2 firing (87 s) was carried out 21 days after the G4-1 firing (53 s), and G4-3 28 days after G4-1 , compared with the 7 day maximum specified . The preparatory operations of filling, pressurisation and countdown were conducted satisfactorily in every case.

CONCLUSION

To sum up, the cluster or G test firings have served to demonstrate the correct behaviour of the Ariane hardware which has been produced by a large number of European firms. The 10 tests have allowed any weak points to be determined and corrected and have allowed adjustments to be made that have resulted in vehicle performance improvements in a number of cases. Consequently, it has been possible to embark with confidence on the series of propulsion-system tests with flight-standard tanks, which should lead to final qualification in 1979.

Stage-Separation Testing for Ariane

An essential element in the development of the Ariane vehicle has been the demonstration of correct functioning of the pyrotechnic devices that will be responsible for first/second and second/ third stage separation during the launch sequence, and satisfactory resistance of these devices to the harsh environment in which they will be called upon to operate. The pyrotechnics for Ariane have been flight-qualified on the basis of both 'component tests' and the testing of complete pyrotechnic chains or 'integration tests'. It is the latter integration tests that form the basis for this particular article.

THE COMPOSITION OF THE PYROTECHNIC CHAINS

The functions to be performed during first/second and second/ third stage separations are identical in as far as both involve: ignition of acceleration rockets on the upper com - posite ignition of retro -rockets on the lower composite and simultaneous cutting of the flange between the stages jettisoning after burnout of the acceleration rockets and their support structures. The sequence of events in the separation chain begins with the sending of an electrical command signal (duplicated) to an 'arming unit', the purpose of which is to convert this current into a detonation, which then travels along confined detonating fuses (CDFs) to the linear shaped charges responsible for structural cutting, and the devices that ignite the separation rockets. These chains are all fitted to the outside of the launcher on structures that also carry other, e.g. hydraulic, pneumatic or electric, functional systems equipment. As the latter equipment is located very close to the pyrotechnic elements, which produce very high-level shocks (150000 g) when the structure is cut, integration testing must assume the double role of: demonstrating on mock-up structures that the pyrotechnic functions are correctly performed ('Category-I'testing) demonstrating, using real structures fully equipped to flight-standard (functional units, systems, etc.). that the pyrotechnic environment will not damage the u nits, systems, etc. ('Category-II' testing) .

THE INTEGRATION TESTS

Four integration tests have been carried out to qualify the separation of Ariane's three stages, and each was planned and executed on the basis of the flow diagram shown in Figure 1. The sequence shown has been applied as a standard procedure in all the major Ariane tests, with the aim of providing maximum visibility as to: exact configuration tested conduct of the test itself results obtained . It can be seen from the diagram that there was provision for a technical review before the test itself, as well as a final check on test procedures, including operating instructions for the test means by which the various parameters were to be measured test result sheets, for both ilnstrument readings and observations. As soon as each test was completed, the first visual observations were noted on the test sheets, the acquisition of measurements verified, and the camera films processed . An indication of the amount of preparation and subsequent effort needed for the separation tests can be gained from the fact that a test sequence lasting only 25 s involved a full seven -hour day by the team, all operations being manual except for the final firing test proper, for which an automatic sequencer was used. A Test Review Board (,Commission de Revue des Essais ') was convened after each test to ensure that the documentation used for the test was that approved or amended during the technical review, that the first results agreed with predictions, and that the measurements acquired would make it possible to process the data successfully and exploit the test results. Provided these criteria had been satisfied, the Board authorised the test set-up to be dismantled.

INTEGRATION TESTING FOR SECOND/THIRD STAGE SEPARATION

The third-stage thrust frame used for these tests had previously undergone the requisite separation and subsystem vibration tests . It was completed by a mock -up skirt representing the second stage in order to accom - modate all the pyrotechnic elements fitted to that stage; only the rockets were dummy units. Both the building up of the fully equipped thrust frame and the test proper were carried out in the main hall of the SNIAS pyrotechnic laboratory at Les Mureaux. A total of 74 parameters were measured during the tests (28 time measurements, 33 shock measurements, 9 movement measurements, 2 pressure measurements and 2 temperature measurements) . Three stages of the separation test are illustrated in Figure 2. The accompanying legend traces the sequence of events from activation of the arming unit until the achievement of stage separation . TEST RESULTS

One of the test objectives was to determine the impulse per metre of the linear shaped charge used for structural cutting . The value recorded did, in fact, considerably exceed specifications, and there proved to be a good measure of agreement between the various tests and the various methods of measurement. As far as the prediction of the trajectories of thrusters and their supporting structures were concerned, the test results agreed well with calculations and it is therefore possible to predict the in-flight trajectories during separation with confidence. Shock -attenuation measurements were also made during the test and a clear correlation was apparent between damping-out time and distance from the linear shaped charge.The levels of shock sustained by various items of thirdstage equipment in the course of the separation are listed by way of example in Figure 3 (duration of the shock shown in brackets) .

CONCLUSIONS

Four successful Ariane stage separation tests were carried out in 1977 and they have demonstrated both proper functioning of the sequences and pyrotechnic devices and satisfactory resistance of the equipment units subject to shock. The scope of the tests - size of structures and volume of information processed - has served to endorse the merits of laying down precise test procedures beforehand and shown just how much information can be obtained from movement and shock measurements of the sort described.

The HM-7 Engine of Ariane's Third Stage

The HM-7 engine, which is the first cryogenic engine developed in Europe, was designed specifically to meet the propulsion requirements of Ariane's third stage. It consists of a thrust unit fed with liquid hydrogen and liquid oxygen under pressure by a turbopump, and devices to monitor, control and condition the propellants. The whole assembly is mounted to the stage thrust frame (Fig. 1) by means of a gimbal and electro-hydraulic actuators, which allow pitch and yaw to be controlled. The main contract for the HM-7 was awarded by CNES to SEP (France), and the subcontract forthe thrust unit to MBB (Germany). The engine's main characteristics are listed in Figure 2 and its operating cycle is shown schematically in Figure 3. It has a conventional gas -generator cycle, in that the turbine gases exit through an exhaust independent of the main nozzle. The liquid hydrogen from the tank enters a centrifugal pump from which it passes to a chamber inlet manifold where some 6% of the flow is used to cool the upstream portion of the nozzle before being ejected level with the nozzle exit, a principle known as 'dump cooling' . The main flow, which is used for combustion, is fed back to the injector throi.Jg h a regenerative circuit made up of slots milled in the chamber body. The latter is cooled by hydrogen circulating in the opposite direction to the combustion gases. The hydrogen exits from the regenerative circuit in gaseous form. A small amount is tapped in order to pressurise the hydrogen tank and to supply the attitude and roll -control system as and when required. The liquid oxygen is injected directly into the chamber as it leaves the pump. A pyrotechnic igniter fires the chamber gases.

The turbine driving the two pumps is fed by a gas generator, which is itself fed by the hydrogen and oxygen tapped at the pump outlets. The gas flow to the generator represents 1 .8% of that to the chamber. The exhaust gases from the turbine produce a slight additional thrust but its specific impulse is less than t hat of the chamber. In order to reduce the possible effects of interacti on between the jets from t he ma in chamber and the turbine exhaust the two streams are kept in the same plane and at similar pressures. A starter cartridge is used for starting the turbine and ig n iti ng the gas generator. Regulation and adjustment The thrust is regulated by the open -loop method: the turbopump's speed is stabilised by adjusting the feeds to the gas generator. The engine torque is kept practically constant w hen the turbine rotation speed exceeds 90% of its nominal value, the oxygen flow to the generator being stabilised by a pressure regulator and a cavitating venturi . A similar venturi in the hydrogen circuit makes the flow of that gas independent of the generator's operation . It was not thought necessary to regulate the mixture ratio, because it can be adjusted to within ± 1 % during acceptance testing by carefully selecting the calibrated orifices immediately upstream of the valves that inject hydrogen and oxygen into the chamber. The engine start-up sequence has three main phases: (i) Before lift-off and during launcher preparation Conditioning of all circuits by flushing with helium and successive compression and expansion. Commencement of cooling of propellant feed lines. (ii) During flight of the first two stages Keeping feed lines cold by venting small quantities of propellant through the flushing circuits. (iii) During third-stage flight On second/third-stage separation, precooling the chamber regenerative circuit for 2.5 s by opening the hydrogen-injection valve and flushing the oxygen injectors with helium. Initiating the igniter. Igniting the starter and opening the oxygen-injection valve. Opening the generator's injection valves.

THE THRUST UNIT

This unit is made up of three subassemblies: an injector, a combustion chamber, and a nozzle. The injector is made from 90 identical coaxial elements equally spaced around five concentric circles. The face plate between injection elements is made of Siperm, a porous material through which a small fraction of the hydrogen flow is able to pass, thus ensuring cooling. Each injection element gives a twist to the liquid-oxygen jet and surrounds it with an annular layer of hydrogen. A high injection-velocity ratio and a 'swirler' ensure good homogeneity of the mixture and help to improve combustion efficiency.

The combustion chamber itself (Fig . 4) is cooled by 100 slots of varying cross-section milled in the copper chamber body and closed on the outside by successive layers of electro-deposited copper and nickel. The configuration of the slots and the thickness of the chamber's inner wall were the subject of complex mathematical modelling to determine the thermal constraints and temperatures in the various areas, particularly in the neighbourhood of the throat where the figures are highest. A detachable nozzle extension (Fig. 5) has been used to adapt the engine to sea-level test conditions by extending the combustion chamber outlet (beyond point where section ratio=7) . It was made from 242 Inconel tubes arranged spirally around a mandrel and argon-welded on the outer periphery. Four hundred and eighty-four small nozzles were brazed into its lower end, enabling part of the impulse from the cooling flow to be recovered. This arrangement has the advantage of separating the nozzle from the combustion chamber when carrying out development tests under sea-level conditions, without danger of jet separation or modification of the engine's operational parameters.

THE TURBOPUMP

A general view of the turbopump is shown in Figure 6. An axial turbine with two pressure stages drives both pumps. Eighty -two percent of the turbine power is used by the hydrogen pump, which is mounted on a common shaft with the turbine, the remainder being transmitted to the oxygen pump through a reduction gear. The turbine rotors are made of I nconel and are machined by electrode erosion . The stators are cast. Each pump wheel comprises an axial inducer, whose vanes are sharpened to improve cavitation performance, and a radial-vane impeller equipped with counter-vanes for balancing the axial thrusts. The generator is of the non-uniform-mixture type. The injector unit is made up of three oxygen injectors converging at 45° surrounded by 12 axial hydrogen injectors.

The starter produces a gas flow at 2000 K for 1 s which first serves to spin up the turbopump and then, after opening the generator's injection valves, to ignite the latter. The propellant-feed circuits have been designed so as to reduce the dead volume between the injectors and the injection valves.

DEVELOPMENT TESTING

The milestones in the development testing of the engine (Fig.7) have been as follows: - July 1973 Start of project - July 1974 First turbopump test - August 1974 First sea-level thrust-unit test - May 1975 First sea-level engine test - November 1976 First long-duration sea-level en - - April 1977 - May 1977 gine test Turbopump qualification First engine test with altitude simulation June 1977 September 1 977 - August 1978 Thrust-unit qualification First long-duration engine test with altitude simulation Start of engine qualification The main test facilities used are listed in Table 1. Test stand PF 41 (Fig . 8) is the largest facility, consisting of two test cells with common services. The steamextraction device, with a flow of 110 kg/s, creates a 50 mbar vacuum to simulate the conditions under which the chamber is pre-cooled and ignited. The thrust-unit test stand (Fig . 9) also comprises two test cells. Altitude simulation is achieved by passing the chamber flow through a supersonic diffuser. The results ofthe tests conducted up until1 May 1978 are summarised in Table 2, including those conducted at propulsion-system level. In view of the large number of tests, a computerised management system for their results has been set up. The data are stored on magnetic tapes and a set of interrogation procedures gives automatic access to the information on the basis of well-defined criteria. This system constitutes a propulsion data bank, which has been used for working out mathematical models for predicting performance and for dynamic studies, as well as for exploiting and summarising test series. Each H M -7 engine undergoes acceptance testing on the basis of the procedure summarised in Figure 10. Initial adjustments are derived from a mathematical model of the engine, and measurements are made to an accuracy of 0.5%.

POSSIBLE PERFORMANCE INCREASE

The margins within which the engine can be adjusted allow its thrust to be increased and its mixture ratio to be altered without any technological modification. The development tests have already allowed the H M-7's upper operating limits to be explored (± 1 0%), and a number of improvements have been proposed with a view to increasing launcher performance. They relate mainly to increased specific impulse, by increasing combustion pressure and expansion ratio, and to the lifetime of the chamber. A new chamber configuration with a reduced throat diameter and increased pressure has already been tested, and it leads to a gain of 2.5 s in specific impulse. A new configuration for the chamber cooling slots has been drawn up to allow increased operating time. It involves decreasing the depths of the slots and increasing their number from 100 to 130. This modification allows the wall temperature to be reduced by 1 00 K and leads to a considerable increase in creep limit.

The Ariane Fairing and its Separation Systems

The essential function of the fairing is to house and protect the payload against such harmful environmental influences as humidity, rain, sunlight, winds and dirt whilst the launch vehicle is on the ground and against aerodynamic loads and heat fluxes during flight. The rather peCUliar, bulb-shaped configuration of this upper part of the launcher is the result of geometric considerations associated with optimum design of the payload bay. To satisfy possible communications needs between payload and ground stations before, during and after launch, Ariane's fairing is specially designed to include a radio-transparent rear cone, and optional radio-transparent doors or areas in its cylindrical or upper elements. Once the fairing's in-flight protective functions have been served, it must be capable of separating from the launcher with a guaranteed clearance with respect to both payload and vehicle. For Ariane, separation will normally take place at an altitude of some 110-140 km, the exact point in the flight path usually being characterised in terms of a particular launcher acceleration (up to 45 m/s2). The general composition of the Ariane fairing (Fig. 1) is as follows. An aft cone with an interface diameter to the equipment bay of 2.6 m opens to 3.2 m maximum diameter and is followed by a cylindrical section 4 m high that terminates in a front-cone section with a spherical nose. The overall height of the fairing is 8.6 m. As can be seen in Figure 2, the fairing can accommodate a maximum payload diameter of 3.0 m and the useful payload volume amounts to approximately 40 m3 . The cylindrical section and front cone are of classical metallic -frame/ stringer construction; the aft cone, for reasons of radio -frequency transparency, is a kevlarglassfibre sandwich . Access to the payload is provided by four 450 x 450 mm doors in the cylindrical section . The positions of these doors can be adjusted w ithin a specified zone to suit the particular payload carried. Access to certain Vehicle Equipment Bay (VEB) items is possible via doors in the rear cone. Standard payload connectors on a variable -length support, to cope w ith a large variety of satellite sizes and connector types, are available. A second such device can be provided for duallaunch purposes. The fairing can be delivered with acoustic protection for those payloads susceptible to lift-off and aerodynamic noise.

DEVElOPMENT PHILOSOPHY

The development philosophy for the Ariane fairings (Fig. 3) foresaw the use of three units for development and qualification on the ground (DMU, SM1 and SM2), and the delivery of four flight units (L01, L02, L03 and L04) . The static and dynamic qualifications were to be realised using two complete structures (models SM1 and SM2) . Three separation tests have been conducted using SM 1 in the large vacuum chamber (DTC) at ESTEC, the fairing unit being refurbished after each test. A linear thrusting joint system is used for fairing separation (vertical), that for the first test being delivered by McDonnell Douglas Astronautics Company (MDAC) . A European -developed system was employed for the second and third tests. A mathematical model to predict fairing dynamic behaviour and separation trajectory has also been developed and it has been applied in separation tests using a fairing rear cone prior to fairing -separation qualification proper. A dispersion analysis allows in-flight fairing -separation trajectories to be predicted taking into account the inherent dispersions of both the fairing and its separation system. The material characteristics of the structure were subjected to manufacturing tests prior to prototype con - struction , A dynamical model (DMU) has been provided for use in overall vehicle checks. This model will also serve for electromagnetic -compatibility and radio -compati - bility checks during electrical mock-up testing. The static qualification unit (SM2) has been refurbished for use during propellant mock -up tests at the Guiana launch site. The electrical system has already been subjected to breadboard testing, and two electrical simulators have been manufactured, also for use in the vehicle electrical mock-up tests. The acoustic protection system will be employed in the fairing to be used on the second test flight (L02) to provide in-flight qualification. Radio-frequency tests have been conducted and a 1/5 th scale-model will be permanently available to verify the suitability of particular antenna configurations. Lanyards, umbilicals, venting and cooling have also been the subject of development and functional qualification testing, the SM2 static model being used in the majority of cases. Fairing development has been entrusted via SNIAS to Contraves, leading representative of the Swiss Aerospace Consortium, which has the Swiss Federal Aircraft Factory, Pilatus and FFA as partners.

+ про системи розділення ГО (ст 80-83)

First Ariane Launcher leaves for French Guiana

A major new phase in the Ariane test programme began in June when the full-scale propellant mock-up (47 m high with a maximum diameter of 3.8 m) left the Launcher Integration Site at Les Mureaux near Paris for the Ariane Launch Base in French Guiana. Three pressurised containers (one for each stage) were transported by barge to Le Havre, where they were loaded onto a freighter bound for Cayenne. Early in July the stages, the fairing and the other elements of the launcher were transported by road from the port of Cayenne to the Ariane Launch Site at the Guiana Space Centre. I n August, the launcher will be erected for the first time on the launch table and the propellant-mock -up tests, which will last for three months, will begin. These tests are designed to check: the general conditions for launcher assembly and compatibility with the vehicle of the ground facilities (platform, tower, etc.); the technical functioning of the fuelling and draining facilities and systems, both on the ground and on the vehicle, and the satisfactory working of the relevant p roced u res. The transport, assembly and test operations will be carried out by Centre National d'Etudes Spatiales (CNES). The system integrator, Aerospatiale, will be responsible for evaluating the dynamic and thermal behaviour of the launcher in- ambient climatic conditions (wind, temperature) and when subjected to vibrations simulating those at lift-off. (11)

Second-stage tests completed Following the propulsion-bay tests carried out in 1977, the next step has been development testing of the complete stage. The purpose of these tests, which have now been completed, has been: to validate the procedures for testing and operating the stage to verify proper behaviour of the structures: tanks, thrust-frame, skirts, water torus to check the flight-standard adjustments of the equipment to check the behaviour of the hardware in the exacting thermal and vibratory environment to verify the operation of the pressurisation system when fed by only three high-pressure helium bottles instead of four as originally planned to simulate a waiting period of the launch pad in Guiana, by heating the propellants to 22.5' C to study the different cases of propellant (UDMH, N2 0 4 ) and water depletion and to evaluate the corresponding thrust decays. It will be recalled that the first static firing (M1) took place on 31 January, and lasted 138 s, the nominal operating time. The M2 stage was tested on 31 March, functioned nominally (138s), and underwent a further test of 17 s on 10 April to study N20 4 depletion. M3 has been fired three times: on 14 August for the nominal duration of 138 s, which enabled simultaneous depletion of both propellants, and then on 18 and 29 August for two short-duration tests (17 and 23s) terminated by the depletion of N20 4 and H20 , respectively. All of these tests, conducted at DFVlR's Hardthausen test centre, have proved very satisfactory, with no incidents of any kind reported. Three other stages are now due to undergo qualification firings, in October (01), in December (02), and in February 1979 (03). Concurrently, tests with the engine adapted to vacuum conditions are continuing, and two qualification tests in vacuo are planned for early 1979. (12)

First-stage development tests The fourth development test of Ariane's first-stage (L 140) in flight configuration was conducted successfully on 5 December at the facilities of Societe Europeenne de Propulsion (SEP) in Vernon, France. The test lasted 142.9 s, a propulsion time corresponding to depletion of the Np 4 propellant, and thus confirmed proper functioning of the propulsion system for the nominal duration of the stage's flight (142.5 s). One of the essential aims was to test the new silicon and phenolicresin based material (Sephen 301) for the throats of the Viking-engine nozzles, which replaces the materials whose resistance was judged to be inadequate on the basis of earlier tests. All propulsion-system elements performed according to specification during the test and, based on an initial assessment, all test objectives were satisfactorily achieved. Third-stage development tests The second test of the third stage in flight configuration, scheduled for 28 November at Vernon, was interrupted following an ignition failure in the cryogenic engine (H8). The slight explosion that occurred damaged the propulsion bay but not the test stand. A new schedule for this stage's testing will be drawn up early in January, but the target of making Ariane operationally available in 1981 should be unaffected. Propellant mock-up The erection of the first stage on 6 December marked the first step in the propellant mock-up campaign at the Guiana Launch Base, the purpose of which is to verify the functioning of the filling, draining and pressurisation systems. The second stage was erected on 11 December, and the complete launcher will have been erected by the end of January. The propellant mock-up tests are due to be completed in early April. (13)

Ariane erected for the first time on the launch table Early in December, a particularly important phase in the development of Ariane began at the Launch Site in Guiana - the propellant mockup tests. Their purpose is to qualify the equipment and procedures for filling the launcher, using a dedicated mockup to represent the flight model in terms of filling, pressurisation, sealing and thermal protection. Real propellants and fluids are used. This exercise comprises three phases: (i) A preliminary phase, intended to qualify the operations for handling the various launcher elements and transporting them from the Launcher Integration Site at Les Mureaux to the Ariane Launch Site at Kourou. It involves transport by barge from Les Mureaux to Le Havre, by ship from Le Havre to Cayenne, and by road from Cayenne to the Launch Site. This qualification took place satisfactorily in June 1978. (ii) A launcher-erection phase and bending-mode tests. These operations took place between 4 December 1978 and 4 February 1979, and allowed verification of: the interface between the first stage and the launch table; the procedures for connecting up the umbilical cables and plaques; and the operations for handling and assembling the payload and fairing. (iii) A launcher-filling phase. The main objectives of this phase are to qualify the procedures, software and ground and on-board hardware needed for the filling, draining, pressurisation and flushing operations, using the real launch countdown and a simulated wait on the launch pad, and to measure the launcher's movements when exposed to the wind and the evolution of the propellant temperatures. A further objective of this phase is to assess the Launch Site's ability to carry out two complete fills of the launcher and five complete fill sequences for the third stage (LOX, LH2 and helium). For these latter two phases, a sequence of operations was adopted that allowed all the objectives to be covered in a series of steps, and the operations planned for real launches to be carried at the appropriate moments. Since less experience is available with cryogenic phenomena, the filling tests on the third stage (H8) will be carried out first, and those on L 140 and L33 will follow only when the procedures for H8 have been perfected. It was originally planned to conduct this test campaign between August and November 1978, but it was delayed by: the fact that it took longer than scheduled to set up the facilities, which are rendered complex by the very nature of the fluids used: liquid hydrogen, liquid oxygen, nitrogen tetroxide, and unsymmetrical dimethyl-hydrazine. the introduction of modifications resulting from the development of the stages. To prevent these delays from compromiSing the overall time schedule for Ariane, the Launch-Site test philosophy has been changed: the phase comprising the validation of the checkout system and the operational electrical checking procedures will be carried out before and not after the propellant mockup operation as originally planned. (14)

The propellant-mock-up operallons were concluded towards the end of May with a simulation of the complete countdown as far as Ho-4 seconds, covering all the operations involving filling with propellants or other fluids. In all, the first two stages were each filled twice and the third stage four times during the whole campaign. The conclusion of this operation represents a major milestone, and the following main conclusions could be drawn: The launch-site facilities worked very satisfactorily, and the modifications to be carried out - generally minorwill not jeopardise the L01 launch date. The automatic programmes for filling and prepanng the launcher were validated follOWing the first automatic fill. ThiS IS a remarkable achievement, bearing in mind the complexity of the operations, particularly for the third stage. The launch teams were able to demonstrate their competence to deal With any problems that arose. In Europe, integration of the L01 vehicle is being completed on schedule. Here again, operations proceeded very smoothly and the very few problems that were encountered may all be considered minor. FollOWing the delays In the acceptance of the flight model of the inertial platform, all launcher integration was carried out with the qualification model. The flight model will be integrated into the equipment bay during July and additional tests will then be carried out to ensure that the system's performances are up to standard. The first qualification firing of the first stage's propulsion system was carried out successfully on 17 May. The second and last test is scheduled for early September. The second qualification test of the Viking engine was nominal in both duration and performance, but subsequent examination of the hardware showed premature wear in a turbopump bearing. Investigations revealed some anomalies in the procedure for assembling the bearings and, although the risk of problems occurring during flight was considered very slight, it was decided to change the beanngs on the flight-model engines of the first and second stages and on the two engines remaining to be qualified. This operation, for which no detailed timetable has yet been drawn up, will be carried out on the stages after integration at the Launch Integration Site in Les Mureaux (SIL). The first qualification test on the new nozzle for the second stage was carried out uSing extreme conditions, with completely nominal results. The test campaign for the 82 propulsion bay of the third stage was terminated after a very good series of six tests, beginning programmes & operations in November 1978, all successful and totalling nearly 3000 s of operalion All the lessons learnt in this campaign have been incorporated into the EP4 propulsion system and will be introduced into the flight stage before shipment to GUlana. The first test of the EP4 propulsion system took place on 6 June. This test was stopped after 7 min when a small hydrogen leak was observed. At that time, all parameters were completely nominal. The hardware tested is in good condition and the problem observed would not have had any effect during the flight of the (empty) third stage. There would appear to be a ground problem associated with the test configuration (stage plus test stand) which does not allow the bay to be correctly ventilated. Remedial action will be taken before the next test. (15)

First stage The first qualification firing of this stage was carried out on 17 May, under nominal conditions; the second and last qualification test is scheduled for early September. Third stage Since the incident with the EP 2 stage during the test in November 1978 there have been: six tests on the B 2 propulsion bay ('battleship'), totalling 3000 s of operation; three tests with a flight configuration stage (EP 4), totalling 1600 s. Ariane launch site The propellant-mock-up operations were completed at the end of May with the simulation of a complete launch countdown to H 0-4 seconds for all the operations relating to filling with propellants and other fluids. Tests on the overall compatibility of the programmes & operations launch site were completed at the end of June with the validation of the ignition and jettisoning phase and a simulated flight. These tests allowed qualification of all the launch installations and facilities. Preliminary flight-readiness review The preliminary flight-readiness review of the L01 launcher, carried out during June, has made it possible to clear the launcher for transport to the launch site in Guiana. Ariane follow-on development The ESA CounCil, on 26 July, approved the implementation of the preparatory programme phase for Ariane follow-on development, which is aimed at increasing the launcher's performance from 1700 kg to 2350 kg in geostationary transfer orbit. Manufacture of additional launchers On 26 July, Council also approved the ordering of long-lead-time components for five additional launchers; this increases the number of operational launchers in production or on order from six to eleven.

Further Ariane Development Tests Successfully Completed Ariane's third-stage propulsion system (EP4) was successfully test-fired for the third time on 23 August on the Vernon test stand of Societe Europeenne de Propulsion (SEP). The test lasted for the sched uled 555 s and cut-off was achieved by a simulated on-board computer command. A similar test lasting 570 s was also conducted successfully on 3 July. With the completion of the three tests, the hydrogen and liquid-oxygen (H8) propulsion system (EP4) has now logged a total running time of nearly 1600 s. The last qualification test of Ariane's first stage took place at SEP on 13 September 1979. The firing lasted for 137 s and an initial evaluation of the results has shown parameters to be nominal. This test, together with the last engine-qualification test carried out on 10 September, concludes the ground qualification of the first stage. The start of the Ariane launch campaign has been scheduled for 1 October, and the first qualification launch is expected to take place between 8 and 18 December.(16)

https://t.me/c/1240576052/288 pochatok +

The existence of these risks means that, unlike an operational launch, achievement of the planned orbit was not the sole criterion of success for the first flight. Admittedly, this would warrant a presumption that the various systems functioned correctly, but It was equally Important to obtain from this first flight the maximum of information on how the launcher functions. The L01 vehicle. like the other three vehicles In the development programme, was fitted with very detailed instrumentalion, enabling more than one thousand parameters to be monitored, and was given very specific obJectives. The information gained from this test is being analysed against very stringent criteria and should enable the operation of each launcher element to be checked In detail The campaign After Integration on the Launcher Integration Site at Les Mureaux, France, the complete vehicle was shipped to Kourou, French Guiana, where it arrived on 27 September. The launch campaign started on 1 October. It was initially planned to last 56 working days, which meant a launch on 15 December. The following are some of the key dates in the campaign: checking of the third-stage countdown (including stage fill) on 21 November, installation of the technological capsule on 3 December, followed by fitting of the fairing on 5 December. On 14 December, the countdown for a launch scheduled for 15 December at 11 .00 local time (14.00 Universal Time) began. Everything went smoothly up to the ignition of the first-stage engines, 2.8 s after which one of the pressure sensors in the first-stage propulsion system flagged the computer commanding the automatic launch sequence of a stage malfunction, and this led to engine cut-off and Interruption of the launch sequence. An immediate evaluation of the telemetry data showed at once that all the engines had functioned nominally, and that the false report of an engine malfunction was due to an overpressure effect in a measurement unit. The interruption of the launch sequence after engine light-up led to an aborted firing, and a number of technical problems had to be solved before operations could be resumed. Thanks to prior contingency arrangements by CNES (Centre National d'Etudes Spatiales) and the firms concerned, reinforcements from Europe allowed the vehicle to be restored in record time to a configuration in which the chronology of launch events could be resumed. In the meantime, steps had been taken to avoid any recurrence of the pressure-sensor phenomenon. The countdown was therefore resumed on 22 December for a launch scheduled for 23 December at 11.00 local time. At H-55 s, again dUring the automatic sequence, a false Signal at the instant of SWitching from ground to on-board electrical power supply, followed by the detection of insufficient pressure in the third-stage helium bottle, led to the launch being postponed once more to the following day.

Finally, on 24 December, at 17 h 14 min 38 s UT, after a countdown interrupted by a number of comparatively minor incidents, the launcher lifted off from its launch table for a totally successful flight, the results of which are summarised below. Results The data below show how the actual orbit, as determined by the down-range station network, compared with the nominal orbit: orbit perigee apogee Incllnallon actual 200.8 km 36021 km 17.5590 nominal 200 km 35753 km 17SO The actual orbit lies well within the permissible dispersion. A preliminary detailed analysis of the parameters recorded during the L01 test flight has been carried out by CNES and the main firms involved in the programme it has confirmed the total success that was apparent from the completion of the flight. The conditions guaranteed to satellites on future operational flights, which constitute launcher-qualification criteria, were already met on this first flight in terms of orbit-injection characteristics: satellite position and velocity, final spin-up, acoustic environment, etc. The dynamic environment alone slightly exceeded specification for 5 s during the secondstage flight, when the expected Pogo effect occurred. It was considered unnecessary to activate the device for correcting this effect dUring thiS phase of the first flight. The funclional characteristics of the three stages, the equipment bay and the fairing agreed exactly with the pre-flight predictions. The propulsion performances of each of the three stages were shown to be appreciably greater than the pre-flight estimates. The gains will be published after the L02 test flight. Comment The first phase of detailed evaluation has confirmed then that the L01 launch was a complete success. From a strictly technical point of view, further comment is almost unnecessary, except perhaps to stress the fact that this success was achieved on the first test flight. Several programme policy decisions merit mention: The decision taken at the start of the programme that, as from the first test flight, the vehicle would be complete and fully activated, without resort to intermediate procedures such as the use of a stage derived from other programmes or the lauflch of partially inert vehicles. In this respect, the success of L01 is a 'first'. The decision not to await complete qualification of the third stage before carrying out the first test flight. The adoption of a conservative technological approach, which has already proved its worth where performance is concerned. Furthermore, from the user angle, this success will enhance Ariane's competitive pOSition, which had already made progress in 1979, following INTELSATs decision to use thiS launcher for its Intelsat-V programme. Suffice It to say here that the L01 launch was followed with extreme interest by everyone in the space business, particularly potential users such as Indonesia, the Arab League, and several American firms operating domestic communications satellites (American Telephone & Telegraph. and Western Union), who are about to embark on new satellite programmes and who must shortly choose their launch systems. Latest situation The high quality of the results from the L01 test flight has allowed the second vehicle to be Integrated without modification, other than the specific activation of the second-stage Pogocorrection system. This second vehicle is currently undergoing integration at Les Mureaux, and it will leave Europe early in March for Guiana, where the campaign will start towards the end of that month with a view to a launch towards the end of May. It will be recalled that this second vehicle will fly two passengers - Firewheel and Oscar - in addition to the Technological Capsule. Firewheel is a German scientific satellite for the study of the upper atmosphere, while Oscar is a communications satellite for radioamateurs, built in Germany for AMSAT Corporation. The last two test flights, L03 and L04, are still scheduled for September and December 1980. In addition to their main objective - to contribute to the qualification of the launcher - they Will be used to launch two payloads of great importance to ESA, namely Meteosat-2 and Marecs-A, as well as an Indian experimental communications satellite (Apple). The experience acquired with L01 has shown that the vehicle's design IS sound and that the launch procedures are altogether correct. The next major milestone IS the flight qualification of the launcher, which In all probability will be achieved early In 1981 , as planned from the outset of the programme. I should like to take this opportunity to extend my heartiest congratulations to the teams of CNES, the European firms and ESA who contributed to the unquestionable success of the L01 launch (17)

L02 launch campaign A detailed evaluation of the parameters recorded during the L01 launch has confirmed the total success of the first flight. The experience acquired during this first launch has made it possible to reduce the length of the L021aunch campaign from 55 to 35 days, chiefly by reorganising the launch teams so that the L02 launch can be carried out between 20 and 30 May. The L02 flight stages left France on 18 March for French Guiana and the launch campaign began on 3 April. Concurrently with the launch activities in Guiana, the final ground qualification tests for the third-stage, qualification of which is required before the L02 launch, are being carried out in Europe. Only a very limited number of modifications have been made to the L02 vehicle compared with L01, namely: activation of the second-stage Pogo correction systems (fitted on L01 but not activated); modifications to the measurements made on the engines of the first stage during the release sequence of the launcher retaining jaws; repositioning of certain measurement sensors. In addition to the Technological Capsule (CAT), the L021auncher will carry two passengers, namely Flrewheel, a SCientific satellite for studYing the upper atmosphere, and Oscar-9, a radioamateur communications satellite.

The First Clients for the Ariane Operational Phase

The past several years have seen many contacts between the Agency and potential users of Ariane, and at the time of the first launch of the vehicle in December last year a number of firm commitments had been made by both European and non-European users. Over the months preceding that first and entirely successfullaunch*, and particularly since, the number of contacts has increased considerably and the Ariane launch manifest is filling rapidly.

The user programme The first user of Ariane beyond the vehicle's development programme, dUring which a number of payloads will be flown, will be the ESA Marecs programme. This programme's first maritime communications spacecraft will be launched on the last Ariane development flight in late 1980, and then a second model will be launched In April/May 1981 on the first Ariane production vehicle (LS). This same flight will also carry the Agency's Sirlo-2 experimental telecommunications spacecraft, and thiS will be the dual launch that will demonstrate the Ariane dual- payload capability. The two Marecs spacecraft represent the first part of the Agency's contribution to the proposed LlC-band INMARSA T maritime communications system and they will establish a high-quality shIP-toshore' and shore-to-shlp operational service. A third spacecraft Will be launched later In order to maintain that system. The first two Marecs launches were committed to Arlane in 1979. The July 1981 launch slot is currently free, having been vacated by the first committed non-European user, INTELSAT, which provides International fixed communications services on a world-wide basIs (initially C-band, and now C and Ku-band). INTELSAT has incurred a number of spacecraft delays in preparing its Intelsat-V series, and has now requested that this first Arlane launch be rescheduled for the end of 1981 INTELSAT has also contracted for two optional launches in April and July 1982, has provisionally reserved the launch slot of December 1982, and In addition proposes to reserve the launch slots of December 1983 and oi February and July 1984, the last three being either for the launch of Intelsat-VIVA re-orders, or for new hybrid spacecraft of a smaller class. The custom of the INTELSAT programme (Fig. 2) is Important to the future of Arlane, and the contractual negotiations currently In hand carry great weight. The October/November 1981 launch slot. currently the second operational launch. is assigned to ESA's Exosat scientific satellite. ThiS mission calls for a highly elliptical near-polar orbit. and In addition requires a fourth stage to be added to the basIc three-stage Ariane launcher. ThiS fourth stage, a derivative of the French Dlamant third-stage motor, IS being developed under the Anane development programme specifically for Exosat. The Exosat mission IS therefore of great Interest. In that It will demonstrate the launch vehicle's near-polar high-energy launch capabilities. The whole programme, Including the fourth-stage development. IS a flxed-pnce undertaking The fourth operational launch Will fly the ECS-1 communications spacecraft, the first of a senes of such satellites to be launched In fulfilment of the Agency's obligations to EUTELSAT to provide an operational fixed-services Ku-band telecommunications system for Europe, for the use of the European postal and telecommunications services. Ariane has accepted the commitment to provide a minimum of four launches for the ECS satellite series at ceiling pnces. through 1986. Generally these will be dual launches with companion operational payloads, although in the case of ECS-1 the proposed partner is a Solar Sail experimental package. As already mentioned, the fifth operational launch in April 1982 is currently reserved for INTELSAT, while the sixth in June 1982 is reserved for satellites belonging to the Western Union Telegraph Company - fixed communications services missions for non-European domestic use at C-band. This latter reservation has been made against a firm price quotation for the launch service, and is under negotiation. The seventh launch, in July/August 1982, is also provisionally reserved for INTELSAT, and there is every probability that the two INTELSAT options for 1982 will be converted to firm launches next month. It is planned that October 1982 will see the first launch of the uprated Ariane, Ariane 3A, which will initially be able to inject some 2330 kg of payload into transfer orbit, compared with the conservative claim of 1700 kg for Ariane-1. The Ariane dual-launch (Sylda) system occupies 140 kg of this gross payload, leaving 2190 kg for the two passengers; by mid1983 this figure will be 2280 kg.

One of the two passengers on the uprated vehicle in October 1982 will be ECS-2; the other will be one of two regional C-band missions that have scheduled reservations in this time frame and which are under negotiation, the exact flight allocations depending upon passenger demands and readiness. There are several such non-European Ariane half-payloads proposed for the period between October 1982 and the end of 1983, including the Arabsat regional C-band mission, various RCA US domestic missions, and the Indonesian Palapa-B system. The ninth and tenth operational launches, in December 1982 and February 1983, will be assigned to INTELSAT and to the French national specialised fixed-service Telecom-1A mission. The latter will be a dual launch, as will that of April 1983, the exact mix of committed users again depending upon user programme readiness.

The twelfth launch, in June 1983, is committed to the US domestic American Telephone and Telegraph (A TT) Telstar-3 mission, for which contract negotiations are quite advanced (as always, on the basis of fixed prices). Still later launches, through operational flight 22 in October 1984, will carry payloads for such missions as the Franco-German TV broadcasting projects, the ESA L-Sat large-platform experimental TV broadcast and specialised services project, the operational Meteosat programme, the French national earth-observation project, for INTELSAT, and for a number of third-party users around the world who are already approaching Ariane.

Conclusion

The trickle of launch-service requests received in 1978/79 threatens to become a flood. Given a successful launch in May (L02), it can be anticipated that many options and early requests will be confirmed. There is already pressure to utilise the small number of reserve slots that have been set aside, and the completion and commissioning of a second launch pad at the Guiana Space Centre (Kourou) by mid-1984 is now a matter of urgency. Meanwhile, the American Delta launch vehicle programme has been extended through 1984; there are indications that Atlas-Centaur production may also be extended. Competition is therefore extremely keen, but Ariane can certainly match the prices of the US expendable launch vehicles and is rapidly gaining the confidence of potential users. When the Shuttle eventually becomes operational it will provide further stiff competition for Ariane launch services, and so Ariane cannot afford to be complacent. Ariane pricing is keen and getting keener, with a follow-on development programme under way that promises to almost halve the unit launch cost. With the Ariane production phase now being reshaped as an industrial undertaking, there are clearly interesting years to come.(18)

Ariane + fourth stage The immediate impact on Exosat of the L02 test-flight failure is the nonavailability of performance figures for the secondstage pogo-suppression device and contamination levels, Some data for the acoustic blankets are expected from the first minute of flight. Ariane's fourth stage motor was successfully test-fired at the Istre test range at the end of May,

Ariane

L021aunch Disappointment was in store for those who attended the Ariane L02 launch on 23 May 1980 at the Guiana Space Centre (Kourou, French Guiana). After a countdown interrupted by minor incidents and a weather-induced holding period, launch finally took place at 14.29.39 (UT). After a normal light-up of the four L 140 engines of the first stage, irregular combustion occurred in engine 0 after the first few seconds of flight The sequence of events was as follows: HO + 3.3 s: Launcher lift-off < HO + 4.4 s: All four engines function nominally up to this instant HO + 4.4 s to HO + 6 s: Chamber pressure in engine 0 begins to fluctuate by ±4 bar, finally oscillating with an amplitude of ± 11 bar at a frequency of more than 1000 Hz. Mean chamber pressure remains nominal. HO + 6 s to HO + 28.3 s: Engine 0 once again nominal HO + 28.3 s to HO + 28.45 s: Recurrence of chamber-pressure oscillation of ± 7 bar in engine D, confirmed on film. HO + 28.45 s to HO + 63.8 s: Pressure in engine D nominal once more. A temperature sensor on the propulsion bay records a linear rise from + 24° to ~C. HO + 63.8 s: Temperature in question rises sharply to 100"C, and the chamber pressure in engine D falls simultaneously to 10 bar. The vehicle experiences a powerful roll torque. HO + 63.8 s to HO + 104 s: The flight- control system succeeds in maintaining the launcher in the nominal trajectory plane. The roll rate reaches 60" per second. HO + 104 s: Fall In chamber pressures in engines A and B, hitherto completely nominal. Engine C continues to function nominally. HO + 108 s: Fall in chamber pressure in engine C and destruction of launcher, probably initiated by the breaking of a structural connection, as a result of the considerable general stresses, activating the self-destruct system fitted to each tank. A number of theories attributing the irregularity either to the engine or to the environment have been put forward. The initial cause of the malfunction of engine o is still to be determined, by thorough study of the various recordings and films available and a comparison with the data collected dunng the first and fully successful Ariane flight, In December 1979. In addition, a search has been undertaken with a view to recovering the wreckage of the launcher from the sea, particularly the first-stage propulsion bay, examination of which might proVide valuable informalion. The L02 failure does not call the contlnualion of the programme Into question. When the cause of the engine failure has been Identified and the necessary corrective actions taken, the programme authorities will conduct the other two planned qualification firings. Six Ariane vehicles are currently being manufactured - within the framework of the launcher promotion senes - for the placing in orbit of the satellites MarecsB/Sirio-2, Intelsat-V, Exosat, ECS and Telecom 1A (see article 'First Clients for the Ariane Operational Phase', ESA Bulletin No. 22, pages 66-69). (19)

Launcher

The Ariane fourth-stage active nutation damper validation tests which are examining potential fuel-sloshing problems associated with Exosat's propellants have been completed and the preliminary results are satisfactory. Assessment of the data available from the early seconds of the L02 flight have led to the conclusion that the acoustic blanket is as effective as predicted.

Cause of the failure of the second Ariane launch (L02) found

The failure of the Ariane L02 launch on 23 May 1980 was due to combustion instability at high frequency (above 2000 Hz) in one of the four first-stage engines 5.75 s after ignition. This extremely violent phenomenon, lasting 0.3 s, abruptly altered the characteristics of the injector, degradation of which led to the destruction of the engine at HO+ 64 s. The fire that then broke out in the propulsion bay caused the vehicle to be destroyed 108 s after lift-off. Much meticulous work was needed to narrow the range of hypotheses for the cause of the failure and finally to reproduce on the test-stand the behaviour of the faulty engine. This involved analysing the telemetry recordings from the launch, inspecting the damaged hardware recovered from the sea, static firings of engines (37 between July and mid-October 1980), acoustic simulation and an investigation of manufacturing and inspection processes. In particular, it has been proved that the cause of the engine failure could not have been external to the engine itself, and the hypothesis of the presence at ignition of a foreign body (e.g. an identification tag or filings) has been eliminated. Furthermore the interaction of acoustic effects between the ground and the vehicle during lift-off was slight, being comparable to that observed during the ground qualification tests. Work by specialists at SEP, ONERA, SNIAS and CNES has led to the conclusion that the high-frequency combustion instability of engine '0' was caused by dispersion in the characteristics of the system for injecting fuel into the combustion chamber. This disperSion probably resulted from slight variations in the manufacture of successive units with respect to certain geometrical characteristics of the injector, the sensitivity of which did not come to light in the numerous engine development tests (nearly 200). In the light of the results attained, it had not been considered necessary to impose stricter manufacturing tolerances. It took a prolonged research effort on some 30 groups of parameters for each injector and correlation with the development tests to reveal the variations in question. (The injector is extremely complex with nearly 1000 injection orifices delivering almost 250 kg of fuel and oxidant per second) In consequence, ESA and CNES have jointly decided to adjust the manufacturing tolerances for the injectors and to select the latter on the basis of static firings on the engine test-stand. Provided that the tests scheduled from now until the end of the year are satisfactory, this programme should allow the Ariane L03 vehicle to be equipped with test-selected injectors for a launch under good technical conditions in the second half of March 1981. The fourth and last test flight would then be scheduled for June 1981, leading to a first operational launch in October, and the subsequent operational programme schedule should be unaffected. (20)

Ariane

APEX programme

In recent months the APEX programme - which allows payloads to be flown on the Arlane programme test flights - has passed two major milestones. Exosat engineering model under test at ESTEC (above) and earlter (right) during integration at MBB (Ottobrunn) Modele d'idenlifleatlon d'Exosat en eours d'essais a I'ESTEC (en haut) et en eours d'lntegratlon ehez MBB, Ottobrunn (a drolte) After preparatory work in Europe had been completed without inCident, the L02 composite payload was shipped to Guiana in early April for the final preparation of the launch. It conSisted of the Ariane technological capsule (CAT), the scientific satellite Firewheel, and the radio-amateur satellite Amsat (Oscar-9). Despite the disappOintment of the unsuccessful launch, the five-week payload preparation exercise nonetheless proved that the teams and facilities at the GUlana Space Centre serving the payloads enable a complex launch campaign to be satisfactorily carried out. The second recent milestone was the final preparation In Europe of the Ariane L03 composite payload, consisting of the CAT, the Indian satellite Apple, and Meteosat-2, which IS designed to take over from ItS predecessor Meteosat-1 . The Indian satellite Apple, which has been deSigned, built and integrated by the Indian Space Research Organisation (ISRO), IS the precursor of the Indian communication satellites of the Insat family. Its miSSion, which IS mainly technological, has the following objectives: to familiarise the Indian space agency With the construction of a threeaXIs stabilised satellite placed In geostatlonary orbit by an Indian-built apogee motor; to acquire experience with the subsystems constituting the L- and Cband telecommunications platform; to train ItS tracking network (ISTRAC) In tracking a satellite from its injection point into transfer orbit to ItS final station, and In subsequent statlonkeeping; and to familiarise personnel with satellite communications, local and long distance, and with the transmission, acquisition and processing of the data. The L03 composite is a major European 'first' because of Its weight (1635 kg) and dimensions (height of 6.525 m measured from the separation plane with the launch vehicle). The Meteosat and Apple satellites having separately and successfully completed their acceptance tests, the assembled flight model of the L03 composite successfully passed its acceptance tests in October and November. The flight-readiness reviews for the two satellites will take place successively in December and January. In parallel with this work, the integration of the Ariane L04 composite payload, consisting of the CAT and the maritime communications satellite Marecs-A, has been proceeding satisfactorily and the qualification tests are scheduled for January at CNES's Toulouse space centre. (21)

Ariane

The injectors for the five engines of the Ariane launch vehicle for the L03 test flight have been successfully acceptance tested on the SEP test stands at Vernon, and are now being fitted. These injectors have undergone the modifications adopted following the investigations and tests carried out in order to overcome the high-frequency phenomena that occurred during the L02 test flight. Acceptance involves two 'hot' tests of each injector under conditions much more severe than those encountered in flight. The results were very satisfactory and, as things now stand, the third Ariane test flight can be scheduled for the second half of June. The Ariane L03 vehicle will fly: a technological capsule (CAT), scheduled for all the development flights, containing electronic equipment and environmental sensors the European meteorological satellite, Meteosat-2, to be shipped to Guyana in mid-April the Apple communications satellite, designed and constructed by the Indian Space Research Organisation (ISRO). The satellite campaign will start in the second week of April and that of the launcher early in May. The launch slot will be fixed at that time. Series production of the launcher has already started. To date, 13 firm orders for launches have been received, together with 12 options or reservations.

Ariane L03 Launch set for 19 June The injectors for the five engines of the Ariane launch vehicle for the L03 test flight have been successfully acceptance tested on the SEP test stands at Vernon, and are now being fitted. These injectors have undergone modifications following the investigations and tests carried out to overcome the high-frequency phenomena that occurred during the L02 test flight. The acceptance has involved two 'hot' tests of each injector under conditions much more severe than those encountered in flight. The results have been very satisfactory and the third Ariane test flight is expected to take place on 19 June 1981 . The Ariane L03 vehicle will fly A technological capsule (CAT), carried on all four development flights, containing electronic monitoring equipment and environmental sensors. The second European meteorological satellite, Meteosat-2. The Apple communications satellite, designed and constructed by the Indian Space Research Organisation (ISRO). (22)

Third Ariane Test Flight Success

After the spectacular success of the first test flight (L01) of the Ariane launcher on 24 December 1979, the failure of the second launch (L02) in May 1980 caused some consternation. Failure of one of the four first-stage engines, owing to combustion instability, resulted in the L02 launcher being destroyed some 100s after lift-off. The investigations and tests to determine the cause of this failure finally led to the conclusion that the combustion-stability margin of the Viking engine, which depends essentially on the geometrical configuration of the propellant injector, was too narrow. A number of static tests were made to determine the most effective modification for the injector, which was then qualified in a series of tests in which the operating conditions were distinctly more stringent than those laid down for the third test flight (L03). The engines for this third flight were equipped with the new injectors, so removing any danger of a recurrence of the phenomenon that led to the failure of the second test flight.

Launch campaign

The launcher and supporting equipment left Le Havre by sea on 17 April and reached Kourou on 29 April 1981 .The launch campaign began on 4 May and lasted for 34 working days (compared with 56 days for L01). No significant diHiculties were encountered. The payload, consisting of the Agency's own Meteosat-2 (the second European meteorological satellite), the Apple* experimental communications satellite being launched for the Indian Space Research Organisation (lSRO), and the CAT (technological capsule flown systematically on each test flight), was erected on the launcher on 3 June. After the fitting of the fairing and final preparations, the launcher and payload were ready for the countdown to start on 17 June.

Launch

The countdown, which lasts 29 h, began at 06.20 UT on 18 June. After two holds in the final phase of the countdown (automatic sequence): the first due to an on-board battery voltage in the third stage being slightly outside limits the second due to one of the two Bretagne radars being out of phase, the launch took place at 12 h 32 m 59 s, (the HO correspondi ng to the moment of the ignition command for the four first-stage engines). During the flight, all the stages and systems of the launcher functioned normally. The separation of the payload composite, of Meteosat and finally of Apple took place exactly as planned, less than 17 min after lift-off, and the satellites were placed in their nominal elliptical transfer orbits. The two satellite control centres - the European Space Operations Centre (ESOC) in charge of Meteosat and the Apple Mission Control Centre in Shar (India) in charge of Apple - reported nominal on-board performances after the first acquisition of telemetry data by the ESA and the Indian tracking networks, respectively.

Preliminary results

The transfer orbit of the composite after separation was well within the nominal tolerances: Perigee altitude Apogee altitude Inclination Actual Planned 201 .5 km 200.0 km 36 175 km 35962 km 10.48" 10.5"

Two modifications made to the L03 launcher since the L01 flight proved their effectiveness: Second-stage Pogo effect It will be remembered that while there was no Pogo effect during the L01 launch in the first stage, the Pogo correction system (SCP) of which had been activated, there was such an effect during part of the flight of the second stage, whose SCP had not been activated. During L03, the secondstage SCP was activated, and this resulted in virtual elimination of the Pogo phenomenon.

PAYLOAD INTEGRATION

Payload contamination During L01 , thermal flux and contamination were observed, both phenomena stemming from the secondstage retro-rockets. The moving of these from the top towards the bottom of the stage, together with the closure of the fairing vents, has solved these two problems.

Conclusion

The launcher's third test flight met the general criteria imposed for a qualification flight and therefore represents an important step towards definitive qualification of the launcher. The excellent results achieved with this third test flight also mean that no major modifications to the L04 vehicle are necessary and the fourth test flight is currently scheduled for November of this year. (23)

Marecs-A to be Launched on Final Ariane Test Flight (L04)

The detailed analysis of the L03 flight data has confirmed the results announced immediately after launch and the total success of the L03 mission (see ESA Bulletin No. 27, pp 75-79) The fourth and final Ariane test flight is scheduled to take place on 18 December 1981 from the Guiana Space Centre in Kourou. On this fourth flight, the European launcher will carry the first (Marecs-A) of two maritime communications satellites which are to be leased to Inmarsat (see elsewhere in this issue), as well as the technological capsule (CAT) designed to provide information on the launcher's performance and which has been carried on all the Ariane test flights. On this fourth flight the CAT will also carry a scientific experiment - Thesee - conceived and built by a team of young people from the GAREF Aerospace Club in Paris. (24)

Successful Launch for Marecs-A

The launch of Marecs-A on 20 December 1981 marked ESA's entry into the international commercial-satellites market, as well as the end of a long period of spacecraft development and testing. Being the first of the series, the Marecs-A spacecraft had been protoflight qualification tested at Sopemea, Toulouse and IABG, Munich, between May and October, the key elements in this test programme being acoustic noise testing, sine vibration testing, solar simulation and full performance verification under extreme thermal/vacuum conditions. The Marecs spacecraft was shipped to the Agency's Kourou launch base on 8 November after being declared flightworthy at the Protoflight Model Review, and a six-week launch campaign was instituted to maintain a pre-Christmas launch. No serious problems hindered progress and on 19 December, at 22.29 h Kourou time (20 December, 01.29 h GMT), after an uninterrupted countdown, Ariane lifted off flawlessly into the night sky. Sixteen minutes later, with spacecraft separation, the Marecs VHF telemetry transponder was automatically switched on and data reception via the Malindi ground station quickly confirmed that everything was nominal on board the satellite. A fine attitude manoeuvre was carried out and the spacecraft's apogee boost motor (ABM) was successfully fired at the fourth transfer-orbit apogee. Soon afterwards, the solar arrays were deployed and the complex sequence of manoeuvres required to achieve three-axis-controlled attitude was carried out. By 22 December Marecs had acquired its nominal operating configuration and was drifting slowly towards its final orbital pOSition at 26°W, which it reached on 2January. Due to the nominal performance of both the Ariane launcher and the ABM, the spacecraft's hydrazine consumption remained very small during this sequence of manoeuvres. Commissioning tests have now been initiated which are aimed at a detailed survey of the spacecraft's health and performance in all operating modes, and at in-orbit verification of the operational procedures. Although these tests are scheduled to continue until the end of January, it can already be confirmed that all Marec's subsystems are performing as expected. Transfer from VHF to C-band telemetry, tracking and control (TTC) is gradually taking place with the handover to the Villafranca station near Madrid. Acceptance testing of Marecs-A will be carried out in February by the International Maritime Satellite Organisation (Inmarsat), which will then lease the Marecs communications capacity throughout the satellite's seven-year lifetime. For those acceptance tests, the Villafranca Payload Test Laboratory (PTL) will be used. In parallel, the second spacecraft MarecsB is being prepared for launch at the end of April 1982 together with Sirio-2, on Ariane vehicle L5. Marecs-B will be placed over the Pacific Ocean area, in accordance with Inmarsat's requirements, and will be operated through the specially built Ibaraki ground station in Japan.

The Ariane L04 Launch Success

This launch, the fourth and last in the development programme for the Ariane-1 version of the launcher, was of the utmost importance because it was required to: confirm the excellent results achieved with the third flight L03, and thus finally vindicate the modifications to the Viking engine's injectors; validate the operational configuration of the electrical systems; complete the launcher-qualification process, which was commenced after the L03 flight. The launch campaign lasted from 3 November to 20 December, and there were no incidents of note, save for the premature release of a LOX valve plate on the third stage which led to a 24-hour postponement, to 01.29 h GMT on 20 December. A preliminary evaluation of the telemetered parameters shows that the 8 launcher functioned nominally and confirms: The correctness of the modifications made to the flight programme in order to achieve greater accuracy at injection; the error at apogee was reduced to less than 10 km. The small impact on the payload of the launcher's thermal and dynamic environments, and the very slight contamination of the payload by the launcher. The smooth operation of the propulsion systems of all three stages. Furthermore, the launcher demonstrated its ability to place a payload of some 1780 kg into transfer orbit. The launch also underlined the operational ability of the launch base, which had to meet very stringent timing requirements, the launch window imposed for Marecs being of the order of only 45 minutes.

Launch Service Contracts

The recent history of space transportation can be considered to have had three phases. During the first, only the USSR and the USA had launchers available. In this period NASA provided a launch service to other western countries whose space technology was less developed, either by furnishing no-cost flights within the framework of bilateral programmes or by providing reimbursable flights for unilateral projects with no NASA interest. In this initial phase NASA showed a certain generosity to its customers in the way in which such aspects as interpretation of reimbursable cost and acceptance of liabilities were reflected in the launch contracts. The second phase began in the seventies, by which time more countries and industries were capable of developing spacecraft, and the requirements for spacecraft for both public and commercial purposes were growing rapidly. At that time the only launchers available commercially were those offered by NASA, the USSR abstaining, except for a few special arrangements, from offering any kind of launch service. It was during this period of monopoly that NASA reinforced its contractual policy as a supplier of launch services, and cost reimbursement was applied much more widely than ever before. Margins were built into the price to cover unforeseen events, and NASA limited its own liabilities to a minimum. Third-party liabilities had to be accepted by the customer, leading to the development of a new insurance market. We are now entering the third phase in the history of space transportation with three launch systems in direct competition: the Space Transportation System (STS/Space Shuttle) conventional, nonrecoverable NASA launchers Europe's Ariane. Other available launchers, particularly those of the USSR, have yet to really enter into commercial competition. The newly competitive environment has caused considerable change in the terms and conditions offered for launch-service contracts. Table 1 shows the contractual arrangements into which the Agency has entered in the past for the provision of launch services. In legal terms these contracts are service contracts, not purchase contracts; they are designed to provide all necessary services to put a spacecraft into orbit as specified in the contract, including any special supporting services. It can be seen from Table 1 that for a number of bilateral projects between ESA and NASA, launch services have been supplied by NASA free of charge. Specific agreements governed the rendering of launch services in these cases and contracts as such were therefore not required. Other ESA scientific projects are covered by the 'Memorandum of Understanding between ESRO and NASA' (MOU) of 30 December 1966, which established the mechanism for obtaining launch services from NASA After an exchange of letters between the NASA Administrator and ESA Director General, to be considered as a political clearance procedure, specific contracts for each spacecraft launch would be agreed. The MOU confirms the principle of reimbursement to NASA of costs incurred and defines the liabilities of both parties. NASA is liable for certain third-party claims and for damage or loss of Government property, within certain limitations. It has been questioned in the past whether experimental (pre-operational) satellites could be considered as falling under this MOU. Although a definitive interpretation has never been agreed, NASA and ESA have always reached agreement in the past on the provision of launch services. A much discussed case 34 was OTS, where there was a question of conflict with the Intelsat agreement. NASA did, however, agree to provide a launch, with OTS being considered an experimental and regional telecommunications satellite. A further problem is whether the MOU could also be applicable for Shuttle flights. NASA has always followed the principle of equal treatment of external customers. Some of the terms and conditions of the MOU are in conflict with the Shuttle-user contract conditions. The MOU could at the most be applied mutatis mutandis.

With the availability of Ariane, the Agency has nominated this European launcher as the preferred carrier for its missions. The first two series of launchings * are under ESA management and no user contracts are required when the Agency uses these launches for its own purposes. In the meantime Arianespace has been created as a private company which will take over the commercialisation of Ariane. All Anane launches beyond the promotional series will be furnished by Arianespace. The basic rules covering the provision of Ariane launch servi:;es to ESA are established in the 'Convention between the European Space Agency and the Arianespace', which applies to Ariane-1. For Ariane-2 and -3 launches, an addendum is presently being negotiated between ESA and Arianespace. Individual launches are covered by specific contracts between Arianespace and ESA, the first one, negotiated for ECS-2, being signed in August 1981. The entry of Ariane has produced a competitiveness in the market and this is now being reflected in the terms and conditions being offered by NASA and by Arianespace.

Pricing policy

In the past NASA's priCing policy on nonrecoverable launchers (Thor-Delta, Atlas-Centaur) was based on the costreimbursement principle. The total cost of a launch contained fixed price elements, such as the purchase of launchers from industry, and cost-reimbursement elements, such as the cost of industrial and governmental support teams as well as overheads. The cost estimates also contained forecast figures for cost-ofliving increases and a margin for unforeseen changes. Although the contracts provided for sufficient retroactive visibility and NASA complied with this in every respect, there were some considerable deficiencies. There was an arbitrary element in the estimates as some important cost elements were not accounted directly against one specific contract, but against batches of contracts. It could happen that, despite an individual launch failure, additional money for incentives had to be paid because a batch of launches showed above nominal performance. There were cost elements completely beyond the customer's control. This method also had the consequence that final invoices normally only arrived several years after the launch, depending upon whether a user happened to be early or late in a batch of launches. In fact NASA is only now submitting the final invoices for ESA's OTS, Meteosat and Geos launches. The Arianespace pricing policy is based on a fixed-price philosophy. The only open element is an inflation factor that is the subject of a contractual price escalation formula based on official indices. To make its launch services more attractive, NASA is now offering fixed prices also. For Shuttle users, NASA offers fixed prices that are valid until 1985, but are subject to revision thereafter. More recently, it has also adopted a fixed-price policy for conventional launchers, margins being included for estimated price escalations.

Launch failures

Conventional launch services are provided on a best-effort basis, so that failures are the risk of the customer, regardless of whether the failure is due to the spacecraft or the launcher. The customer seeking a launch is therefore well advised to assess the implications carefully. He should first establish whether and at what time a back-up spacecraft would be available. NASA contracts on conventional launchers normally foresaw replacement launches being made as and when a slot was next available, with no guarantee of a specific date. If a backup spacecraft is available, the customer should therefore consider booking a back-up launch. He risks losing money if the back-up is not required, but the likelihood of recovering his investment through an alternative customer being found is extremely high. launch service contracts For the Shuttle, NASA has adopted a different policy in that it guarantees one re-flight at no additional charge if, through no fault of the user, the prescribed orbit is not achieved. This is another notable deviation from their former 'no loss, no profit' policy. The procedure in the event of an Ariane launch failure is presently covered by the Arianespace - ESA convention for ESA launches. Because Ariane has been financed by the ESA Member States ESA has been able to obtain special conditions for back-up launches. Somewhat similar to the Shuttle-user conditions, a distinction is made between launch failures due to the launcher and those due to the spacecraft. In case of an Ariane failure, ESA would have the opportunity of obtaining the first launch slot compatible with the availability of the back-up spacecraft. In the case of a spacecraft failure, Arianespace would do its best to assign in the first slot, and in any event a slot not later than 10 months after receipt of a written request. Arianespace generally quotes fixed prices for back-up launches. The Agency takes the view that for its launches the prices agreed in the Convention shall apply. Additional costs due to the accelerated availability of a launch shall be reimbursed by ESA as incurred. There might of course be customers who would already prefer fixed charges, but it is felt that it would be premature to try to establish realistic fixed prices now, because of the lack of relevant statistical data. Another aspect is the coverage of financial loss (cost of back-up launching, loss of income) by insurance. The Agency has taken out such insurances for specific projects, and this aspect has been covered in some detail in Bulletin No. 16.* (25)

Ariane Double Launches - An Operational Challenge

The next Ariane launch will carry two spacecraft simultaneously into orbit. The European Space Operations Centre (ESOC) and its associated ground stations will be called upon for the first time to give full support to two spacecraft in the launch and early orbit phase simuhaneously. The challenges of supporting this double launch are reflected in the flight-dynamics operations, which are outlin·ed in this article.

Ariane's ability to launch two spacecraft simultaneously and to inject them both into geostationary transfer orbit has already been proven by the combined launch in June 1981 of ESA's Meteosat-2 spacecraft and Apple, the Indian communications test satellite. On this occasion ESOC was called upon to give only limited support to the Indian spacecraft. For the next launch, ESOC will be responsible for fully supporting both satellites : the Agency's second maritime satellite Marecs-8, to be positioned over the Pacific and the Agency's Sirio-2, a laser-time-synchronisation and meteorological-data distribution satellite to be positioned over the Atlantic. The two spacecraft missions wi ll share the same ground-station network and both will be controlled from ESOC. Present planning requires that Sirio-2 be injected into near-synchronous orbit at its second or fourth apogee pass in transfer orbit while for Marecs-8 this injection into synchronous orbit is planned for the third apogee. Subsequently, both spacecraft will drift to their operationallongitudes25° West and 177° East, respectively - where actual payload operations will begin. In highlighting the particular challenges in flight dynamics encountered in the preparations for and operations during a double launch, this article concentrates on the orbital and attitude characteristics of such a double launch, including the critical spacecraft manoeuvres needed to ensure correct injection of both spacecraft into geostationary orbit

The launch and early-orbit phase

Figure 1 illustrates the geometrical principles underlying a geostationaryorbit injection procedure. After Ariane's lift-off from the ESA launch site in Kourou, French Guyana, and after a powered flight of approximately 15 minutes, the launch vehicle's third stage and the two spacecraft enter a transfer orbit with a period of 10.5 h, a perigee of 200 km and an apogee of 36000 km. This apogee corresponds approximately to the radius of the circular orbit that a spacecraft must have in order to appear stationary from the ground. Once in transfer orbit, the launcher's third stage performs a pre-programmed attitude re-orientation. It is then spun-up to about 10 rpm, before the two spacecraft are released. Spring forces cause the three bodies to enter slightly different orbit trajectories, reducing the risk of subsequent collision. A 0.5 m/s separation velocity causes the orbits of the two spacecraft to differ by 30 km in apogee and 35 s in orbital period. When looking at Figure 1, one must bear in mind that the plane of the circular, geosynchronous orbit coincides approximately with the Earth's equatorial plane, while the elliptical transfer orbit is inclined some 100 to the Earth's equator. Approximately at the intersection of the two orbits, near apogee, the spacecraft apogee motor is fired to roughly double the orbital velocity so that a near-circular orbit results. The exact timing and direction of this firing depend on the characteristics of the transfer orbit. They are determined by a sophisticated mathematical optimisation which takes into account the total propellant on board and the characteristics of the intended geosynchronous orbit. Although the spinaxis direction, which is aligned with the apogee motor, has already been set by the third-stage attitude re-orientation before separation, the final optimal attitudes can differ by the order of 10" and further reorientations using the spacecraft's own propulsion system are therefore needed during transfer orbit. In view of the load on the ground facilities and the various activities necessary to prepare for apogee-motor firing, it is highly desirable to perform the burns for Marecs and Sirio-2 at different apogees. Once the two spacecraft have been put into near-circular orbits, changes in their orbital characteristics will be introduced so that they will drift, over a period of about 20 days, to their operational locations, where they will be stopped and will remain on station for their operational lifetimes.

The ground-support system

The ground-support system and its functions are shown schematically in Figure 2. A key role is played by the Control Centre, where all data-networking and spacecraft operations are centralised, including all flight-dynamics activities. Most hardware and software facilities for the support of these functions are available in fully redundant form to guarantee operational availability. Linked to the Control Centre are four ground stations, shown in Figure 3, which provide direct contact with the spacecraft and make it possible to receive both spacecraft telemetry streams in real time. Telecommands to the two spacecraft and the necessary tracking information are also handled by these stations. They provide continuous contact with both spacecraft for the first six revolutions in the transfer orbit, except for half-hourly gaps at each perigee. Orbits numbers one, four and six have simultaneous visibility from two or more ground stations, whereas the other revolutions have only single, or partial dual station visibility. The overlayed ground track in Figure 3 gives an impression of spacecraft orbital movements until apogee 4. The Agency's ground-support system for the launch and early-orbit phase was originally established to support only one spacecraft per launch. With the introduction of the Ariane double-launch concept, modifications had to be implemented, such as equipping the ground stations with dual downlink capabilities. Other facilities are shared between the two spacecraft by planning operations such that major activities do not take place Simultaneously. Typically, the ground-station uplink, some of the Control-Centre computers and the flightdynamics operations are largely timeshared between the two spacecrafUn the spacecraft-operations area, dedicated teams are available for each project, while in the flight-dynamics area one team will handle both missions. For the latter, a total of 22 specialists will support the launch and early-orbit phase around the clock.

Flight-dynamics activities

The planning for a double launch calls for detailed time-lining for the flight-dynamics activities to be performed during transfer orbit. The station-acquisition sequence in the near-synChronous orbit is not as compressed as the transfer-orbit sequence and does not require such detailed time-lining.

Once the appropriate sequence has been set up, the overall ground system has to be validated. In the case of the flightdynamiCS activities this will be done via Double Launch System Tests for validation of: the operational computer software the spacecraft and mission parameters the events sequence (nominal and non-nominal) vis-a-vis the computer resources and manning schedule. The tests have to be conducted in the real operational environment. Only in this way can readiness for launch be demonstrated and a guarantee given that the demands of nominal, and more importantly non-nominal, situations can be coped with. The major flight-dynamics operations to be conducted during the transfer-orbit and near-synchronous-orbit phases are: orbit determination attitude determination near-real-time monitoring manoeuvre preparation system identification quality control computer and data interfacing. Orbit determination is performed by iterated least-squares fitting of a numerically integrated arc of an orbit to one or several hours of accumulated tracking data. After determination, the orbit is extrapolated forward in time for planning purposes. Before apogee-motor firing, the transfer orbit that has been achieved must be determined to within a few hundred metres in apogee and a few seconds along track. Attitude determination is performed in a similar way, processing accumulated data at the end of a coverage period with additional computer runs early in the transfer orbit and just before apogeemotor firing. Attitude determination is performed from Sun and Earth-sensor data in the spacecraft telemetry. Depending on the spacecraft's attitude and sensor mounting angles, its infrared pencil-beam Earth sensors have Earth coverage for about 2- 3 h around the apogee of the transfer orbit. Figure 4 shows, as an example, the beginning of the +6" pencilbeam sensor coverage for Sirio-2. Depending on the Sun-Earth geometry and the quality of sensor calibration, a final attitude accuracy of 0.5°-1° can be achieved. Near-real-time monitoring is a determination of the instantaneous attitude and/or spin rate from each telemetry format as soon as it arrives at ESOC. It is particularly useful during manoeuvres, allowing the spacecraft operator to abort a manoeuvre in the case of an anomaly. Manoeuvre preparation has an optimisation and a telecommandgeneration aspect. The optimisation establishes an optimal sequence of attitude and orbit manoeuvres, such that the mission objectives are achieved with minimum fuel. The telecommand generation models the spacecraft dynamics and electronics for thruster firing. System identification concerns the integrated identification of the mathematical-model parameters: environmental parameters, and spacecraft-actuator (thruster) and sensor characteristics.

Quality control supports the planning and preparation phase to establish confidence in the reliability of the overall flight-dynamics support. In addition, selected control activities are performed during the flight. Computer and data interfacing provides the means for smooth operations in the computer environment. Sequence of events The transfer-orbit planning must consider the.requirements of each mission and any constraints on the ground operations, yet it must be sufficiently flexible to accommodate variations in Earth-sensor coverage times and in other parameters. As there are two project groups involved, which naturally have conflicting interests and priorities, it is necessary to set up and plan the sequence of events as a compromise between the various interests and constraints. One of these, the time-sharing of resources, leads to a few general principles that must be applied in double-launch support: manoeuvres shall be carried out with only one spacecraft at a time; the apogee motors of the two spacecraft shall be fired at different apogees, since this is the most demanding part of the ol=lerations; the flight-dynamics operations - orbit and attitude determination and manoeuvre preparation and execution - are pertormed exclusively for one spacecraft during half the orbital revolution preceding its apogee-boost-motor firing. Bearing these prinCiples in mind, one has to prepare an early-orbit-phase operations time line such that the sequence, orbit and attitude determination followed by manoeuvre preparation, execution and system identification is iterated for each spacecraft throughout the transfer orbit, to provide the best possible conditions for the apogee-motor firing. This control loop is shown schematically in Figure 5. The processing of the tracking data for orbit determination is normally not timecritical. It is convenient to process several hours of tracking at times when no manoeuvres or other demanding activities are required. After each major determination, the target direction of apogee-motor firing is updated. In the first few hours of the transfer orbit, however, there is great interest in establishing the mission's success and measuring the orbit for inputs to the attitudedetermination and manoeuvreoptimisation programs. There is an analogous requirement for orbit measurement a few hours after apogeemotor firing. It is also desirable for reasons of tracking geometry to perform a last orbit determination before apogee-motor firing with tracking data obtained from the current station pass. Attitude determination utilises at least 20 min of accumulated Earth-sensor data. It is therefore performed during or immediately after the Earth-sensor coverage periods. The results are needed to prepare attitude slew manoeuvres and to confirm correct spin-axis pointing for apogee-motor firing. The critical events in the entire sequence are the attitude and orbit manoeuvres. The attitude slew manoeuvres are preferably performed within the Earth-sensor coverage periods. The necessary telecommands must be available for uplinking 30 min bef9re manoeuvre execution. Figure 6 shows the presently planned sequences of events for Marecs-B and Sirio-2. The events for Marecs-B are: a crude spin-up to 60 rpm, an attitude slew manoeuvre, a fine spin-up to 65.5 rpm, two further slew manoeuvres and apogee-motor firing at the third apogee pass (A3), followed by three spin-down manoeuvres for the three-axis acquisition. The manoeuvres to be executed by Sirio-2 are: a spin-up to approximately 90 rpm, three attitude slew manoeuvres, and firing of the apogee motor at the fourth apogee pass (M). The operational challenge The planned sequence of events can not be made too tight, because experience shows that time margins are needed to absorb delays in operations. In a double launch, delays caused by the operations of one spacecraft may influence the timing of events for the other and a priority schedule must be agreed beforehand to resolve any conflicts that develop. Changes in event timing are often necessary even during nominal operations because: the orbital period depends on the transfer orbit achieved; the optimal time for apogee-motor firing is not exactly at apogee, but may be up to half an hour before orl after, depending on launch time and measured transfer-orbit parameters; the Earth-sensor coverage times depend on the spacecraft's attitude. The optimal attitude for apogeemotor firing depends on the launch time and measured transfer-orbit parameters. Various intermediate attitudes result from the different slew manoeuvres; the number of slew manoeuvres needed to arrive at the optimal apogee-motor firing attitude depends on the performance of the on-board control system and on the quality of the intermediate attitude and orbit determinations; the attitude determination may need more sensor coverage data than expected, due to Earth and Sun nearcolinearity or Earth-sensor blinding by the Sun. Delays in events can also occur because of the following relatively harmless contingencies: an extra slew manoeuvre may be needed soon after launcher separation if the initial solar-aspect angle is out of limits; the start of a nominal slew manoeuvre may be delayed due to ground-station or on-board system problems; bad-quality tracking data or tracking calibration may necessitate more tracking measurements and a delay in accurate orbit determination; low-quality sensor data or sensor calibration may delay accurate attitude assessment until more data has been collected; a temporary failure in a vital component in the computer system can delay operations until a repair or reconfiguration can be effected. More severe contingencies caused by serious malfunctioning of the spacecraft or launcher can demand considerable efforts at the Control Centre in order to save the mission or to define a meaningful alternative. Such situations can only be prepared for in a general manner by providing the Control Centre with adequate expertise and computer systems capable of performing 'real-time mission analysis'. The flight-dynamics team's challenge is one of preparing a realistic sequence of mission events, implementing it successfully, and being capable of dealing with both planned and unplanned events quickly and competently.

Conclusion

The forthcoming double launch will be a unique event for ESA, and a stepping stone towards future launch-support requirements associated with progressively more powerful Ariane launchers, At the time of writing (April 1982), the miriad preparations at ESOC for supporting this first launch of two ESA spacecraft have already begun. Experience gained in the support of six previous single-spacecraft launchesincluding the now famous Geos-1 rescue - provides a high level of confidence that the success rate can be maintained, espeCially with the availability of standard support facilities that have already been well tested operationally, in the hands of well-qualified and experienced specialists.

Ariane Qualified and Fully Operational

On 25 January, following the complete success of three of its four test flights, the Ariane launcher was unanimously declared flight-qualified by representatives of the European States participating in the Ariane programme. These representatives, from Belgium, Denmark, France, Germany, Italy, The Netherlands, Spain, Switzerland, Sweden and the United Kingdom, took this opportunity to congratulate all those in the Agency, the Centre national d'Etudes spatiales (CNES) and European industry who have contributed to this impressive achievement. The results of the flight tests have confirmed that the Ariane launcher has attained a degree of operational readiness that compares very favourably with that of other available launchers. The declaration of qualification brings to a close the development phase of the Ariane launcher programme, and has firmly established Europe's independent launch capability. In the ensuing operational phase, the first series of seven launches, known as the promotion series', will be carried out under ESA auspices. Thereafter, responsibility for the marketing, manufacture and launching of Ariane will be handed over to Arianespace. The accompanying table is a manifest of the Ariane flights already reserved by international, national and commercial organisations. The length and multinational nature of the list of clients who have already committed to Ariane is regarded by the Agency as a particularly gratifying vote of confidence in Ariane-1 and its successors. (26)

The Ariane vehicle

The initial choice of launcher (Delta 2914) was reconsidered during the project definition phase in 1977, when feasibility studies demonstrated that the European Ariane, augmented by a fourth stage, would satisfy Exosat mission requirements. Ariane's northerly ascent trajectory from the launch range in French Guiana (Fig. 7), and the lack of a restart capability on Ariane's third-stage motor, necessitate the use of an additional stage to inject the 510 kg satellite into orbit. The standard Ariane launcher (Fig. 8) has been uprated by adding a solidpropellant motor (P07). Aside from the motor and satellite adapter, this fourth stage includes a timer or sequencer, an active nutation damper, spin-up nozzles, a despin system, and a telemetry package for transmission to ground of essential performance parameters. After orientation of the composite (satellite + P07 stage) in space by the third stage's attitude and roll-control system for orbit injection at perigee, and upon completion of the extended coast phase, the main fourth-stage events are: initiation of timer by third-stage guidance computer; separation of composite from third stage; spin-up by four nozzles to 47 rpm ± 3 rpm, to achieve spin axis stability during the 2.5 h coast phase; active nutation damping by redundant pneumatic thrusters to inhibit any increase in nutation angle due to external disturbing forces or internal energy dissipation (fuelsloshing); de-activation of nutation damper; fourth-stage ignition; despin after fourth-stage burnout by means of the yo-yo system; separation of satellite several seconds after burnout; depointmg of stage by means of a turnover system. After injection into orbit, some 9845 s after lift-off, the satellite will perform a number of autonomous operations, such as despinning and Sun acquisition. Once Sun acquisition has been achieved, the antenna booms will be deployed automatically. Ariane launcher development was satisfactorily concluded with the last trial launch (L04) in December 1981 and the launcher thereby achieved qualification in standard configuration. Fourth-stage qualification tests are in progress and are expected to be completed in time for a November launch.

The ground segment The ground-segment configuration for Exosat is shown in Figure 9, in which the interfaces within the ESOC Operations Control Centre (OCC) and between ESOC and the supporting station(s) are identified. The supporting stations are responsible for the classical functions of: Satellite telemetry reception at the dedicated ground station and transmission to the OCC where processing and conversion to engineering units facilitates real-time monitoring of satellite status, particularly attitude determination and attitude and/or orbit change manoeuvres. Filing a(ld archiving of technological and scientific data offers the prime users, namely the Exosi;lt scientists, retrieval of up to 24 h of the most recently filed data for further use in the Dedicated Control Room (OCR), Telecommand transmission and verification through the prime station (Villafranca) or any other station required to assist in controlling the subsystems in real time, reloading the OBC to maintain/update the operational status, and supporting orbit changes for occultation and arbitrary pointing. Ranging, a largely automated radiomeasurement function providing: ranging data from the ground station to the OCC's orbitdetermination computer; orbital elements for precise prediction of satellite position; predictions for station coverage to enable the spacecraft controllers to schedule operations. The Observatory Team has at its disposal the Dedicated Control Room, with direct access to the relevant scientific, technological (housekeeping) and operational (mission planning) data. Apart from the real-time data-assessment facilities, data processing is performed off line, resulting ultimately in data tapes for final processing and analysis at participating institutes. The satellite will be 'handed-over' to the Operations Manager at ESOC by the Project Manager, once the scientific payload and satellite subsystems have been checked out in orbit. CSG operations planning Launch operations will begin immediately after satisfactory completion of the Flight Readiness Review (FRR). To ensure effective and efficient preparatory activities at the launch range, meticulous care and attention have been paid to the detailed definition and specification of the activities to be performed prior to the actual launch. This is reflected in the summary launch-operations plan for a nominal launch campaign of six calendar weeks, preceded by a limited Exosat experiment preparation phase of approximately one calendar week at the launch site (Fig. 10). The launch operations are essentially divisible into three main phases, conducted at three different locations on the range. Phase 1 A functional performance test to demonstrate the integrity of the flight model after transport from Europe to CSG, will be performed in Building S1 . Measurements will be made to verify alignment parameters after transport. Similarly, an absolute-leak-rate measurement will be made on the reaction control equipment (RCE) propellant system to ensure that no degradation has occurred as a result of handling and transport. The data resulting from these vital tests are to serve as reference data for the subsequent orbital phase. Final checks on pyrotechnics circuitry and adjustments to mechanisms/appendages, partial installation of superinsulation, and installation of pyrotechnics will take place during this phase, for reasons of internal accessibility and prerequisite to the hazardous operations in Phase 2. Phase 2 The filling of the RCE propellant systems with hydrazine and propane, will be performed in the 'hazardous zone' in Building S3. The experiment gas flushing and replenishment system will be filled and the RCE hydrazine system pressurised in Building S3. Further completion of external multilayer insulation (MU) after battery installation, solar-array installation, propellant loading and associated visual inspections will achieve the readiness status compatible with the milestone for transportation of the satelfite to the launch tower for mating with the Ariane launch vehicle. This phase is completed with transportation of the flight model to the launch tower in the Ariane payload transport container. Phase 3 The activities performed in this phase are constituent elements of the combined operations plan (POC) produced jointly by the Exosat Project and Launcher Authority. The main items are: satellite mating to launcher fourth stage (PO?) satellite checkout after mating satellite monitoring and battery charging flight configuration completion, Le. removal of protective covers, completion of superinsulation and inspection countdown dress rehearsal. After completion of the flight configuration, the flight-satellite arming will be conducted just prior to the start of the composite terminal countdown. The planned duration of this terminal countdown, during which all supporting elements (Le. launcher composite, range and ground network) are activated, is 28h.

Acknowledgement

The Exosat programme has presented a major challenge to all directly engaged in the project and can be considered a major step forward in European satellite development. The authors would like to take this opportunity to acknowledge the often considerable efforts made by the industrial contractors' and ESA staff in supporting the Exosat programme.(27)

Ariane L5 Launch Investigation

Initial investigations into the failure of the third stage of Ariane during the L5 launch on 10 September, suggest that it was due to a problem with that stage's turbopump. The launcher's first and second stages performed nominally, and the third-stage engine ignited correctly 285 s into the flight. It functioned correctly for 275 s, at which point the turbopump speed dropped by 1000 rev/min; 1 s later its speed fell from 60000 to 30000 rev/min, and the pump finally stopped after 325 s of running. The drop in thrust and the consequent engine cut-out caused the loss of the launcher. The most probable reason for the cut-out is thought to be damage to the reduction gear trains in the turbopump, owing to either: faulty lubrication during calibration and acceptance and qualification testing on the ground, or a defect in the gear trains themselves (this hypothesis remains to be validated or rejected by the Board of Enquiry mentioned below). To ensure that no such failure can recur during the next launch (L6). ESA and CNES have set up: a Board of Enquiry composed of European experts, whose task is to determine the cause of the malfunction. This Board is due to submit its results in mid-October an Ariane Programme Review Group, whose task is to check that the L6 launcher conforms to qualification and manufacturing specifications prior to authorising its shipment to the launch base in French Guiana. Ariane's guidance and flight-control systems, and stage and fairing separation systems performed flawlessly during the L5 flight. (28)

Launcher

Following the decision to launch ECS/Amsat on Ariane L6 in the second half of April 1983, as noted above it is now foreseen to launch Exosat on Ariane L7 some two months later, towards the end of June. It is assumed that the panel reviewing launcher aspects and interfaces to satellite hardware and operations within the framework of the Exosat Flight-- Readiness Review can complete its work by the end of December.

Ariane

Findings of the Board of Enquiry and the Programme Review Group On 10 September 1982, the two European satellites Marecs-B and Sirio-2, which constituted the Ariane L5 payload, failed to be placed in orbit because of a malfunction of the launcher 9 min 20 s after lift-off. An analysis of the telemetry and radar data carried out immediately after the malfunction indicated that the latter was probably due to a breakdown in the turbopump of the third stage engine. This initial hypothesis was confirmed by the Board of Enquiry set up by the Directors General of the European Space Agency and the Centre National d'Etudes Spatiales to determine the causes of this failure. In its report, submitted on 15 October 1982, the Board of Enquiry states that, after examining all the available data, it forthwith concentrated on the turbopump. After studying the various possible causes of failure, it adopted as the most likely hypothesis damage to the turbopump gearing due to a combination of the following factors: insufficient lubrication of the gearing during the ground testing of the thirdstage engine before its integration into L5; an unduly narrow operating-safety margin for the gearing due to an unfavourable combination of the various tolerances which, taken individually, were all within the specified manufacturing limits. In the absence of specific telemetry data, an interruption of lubrication in the course of the flight could not be definitively ruled out, but seemed unlikely. In the light of these conclusions, the Board made a number of technical recommendations for improving the performance of the gearing by specifying more precisely all the operations carried out on it. Two types of recommendation were made: short-term measures comprising validation of the acceptance procedure and tightening up of quality control medium-term measures aimed at a new definition for the gearing. It has accordingly been decided to dismantle the turbopumps already manufactured for the subsequent launches in order to implement these recommendations. While the Board of Enquiry was at work, the Agency and CNES undertook a review of the Ariane programme. The Programme Review Group recommended: in the short term, debugging of the hardware for the L6 and L7 launchers in the medium term, essentially tightening up the technical management methods, with particular reference to quality control in the manufacturing processes. The additional work resulting from the short-term recommendations leads to the launch of the ECS-1 and Amsat satellites by the L6 vehicle in the second quarter of 1983, subject to the findings of the L6 Flight-Readiness Review, to be held in January 1983 for the launcher less the third-stage, and in March for the thirdstage alone. To minimise the impact on the launch schedule, it is planned to despatch the third stage by air and to speed up the launch rate by reinforcing the launch team in order to allow a launch to take place every two months. (29)

Ariane Technical status of the programme

The technical recommendations of the Board of Enquiry, which related to the quality of the turbopump gearing and the operation of the lubrication system (see ESA Bulletin No. 33 p. 53), have been implemented. Versions of the turbopump gearing and the lubrication system from which these defects have been eliminated have been developed and tested. All the work involved is being submitted to a detailed review process aimed at confirming the flight-readiness of the L6 launcher. After running-in and inspection, the L6 turbopump has been fitted to the thirdstage engine, due to undergo hot acceptance testing by mid-March. This will be followed by assembly of the propulsion system and then of the complete third stage, which will be dispatched to Guiana in late April. The first and second stages will be shipped in March. Concurrently with the specific action on the turbopump, certain important launcher elements have been intensively reviewed to improve reliability: namely, the inertial-platform system, and the thirdstage feed and pressurisation systems.

Launch schedule

In the light of the foregoing, at its meeting on 23 and 24 February, the ESA Council confirmed its unanimous confidence in and support for the Ariane programme and adopted the following launch schedule: the Ariane L6 launch is scheduled for Friday, 3 June 1983, and the L7, L8 and L9 launches for 26 August 1983,4 November 1983 and January 1984, respectively. Mindful of the interests of the Agency's programmes and of the need to preserve the confidence shown by other Ariane customers and of the time-schedule commitments made, the Council has taken the following steps to ensure that all payloads are launched as soon as possible: To reproduce a mission profile resembling that of the L5 mission as closely as possible - injection of two satellites into geostationary transfer orbit by means of the Sylda dual launch system - Ariane L6 will launch both the ECS-1 and Amsat satellites. For Exosat, the scheduled L7 launch date provides insufficient safety margin vis-a.-vis the closing of the launch window and, moreover, there is a risk of certain experiments in the payload deteriorating through storage. Consequently it has been decided to use a Thor-Delta 3914 launcher to place Exosat in orbit, the launch to take place from Vandenberg in late May 1983. The Ariane-1 vehicle remaining available at the end of the promotion series will be assigned to the launch of Giotto, in July 1985. The L7, L8 and L9 launchers have been assigned to the Intelsat-V satellites F7, F8 and F9. The first Ariane-3launch (L 10) is currently scheduled for March 1984. The Agency's satellites ECS-2 and Marecs-B2, the French satellites Telecom-1 A and B, the Arab League satellite, Arabsat-1 , as well as the American satellites Western Union's Westar-6, Southern Pacific's Spacenet-1 and 2 and GTE's G-Star 1 and 2, will be launched by Ariane-3, the more powerful version of the launcher, capable of injecting two 1195 kg spacecraft into geostationary transfer orbit. (30)

The Ariane L6 Dual Launch: an Unqualified Success

ESA's ECS-1 spacecraft and the Amsat P 3 I b radio-amateur satellite were launched successfully by Ariane (L6) on 16 June at 11 h 59 m 00 s (Ho: ignition of first-stage engines; lift-off at Ho + 3.8 s). The launcher stages and its various systems performed perfectly. The ECS-1 spacecraft was separated from the launcher at Ho + 954.4 s and the Amsat spacecraft at Ho + 1072.4 s. Separation conditions (attitude and spin rate) were as planned for ECS-1, though a perturbation, the origin of which has not yet been determined, occurred after the separation of Amsat. Propulsion from the launcher's three stages was nominal; in particular no Pogo occurred during second-stage flight and the third-stage powered flight was very quiet. Stagings and fairing jettison also went smoothly. The geosynchronous transfer orbit achieved was very close to the nominal orbit: Perigee Apogee Inclination 198.7 km (for 200 km required) 35790 km (for 35870 km required) 8.605° (for 8.60~ required) The launch campaign The L6 customer spacecraft, ECS-1 and Amsat, arrived in Cayenne (French Guiana) on 14 April, and their preparation for launch started on 15 April. Spacecraft integration and test activities in the S1 building were completed on 26 April for Amsat, and on 11 May for ECS-1. 8 Final preparation for launch of both spacecraft was carried out in the S3 building, where ECS-1 was fuelled with hydrazine and its apogee boost motor (Mage-2) fitted and where the Amsat propellants to be used with a Symphonie type engine were loaded. The spacecraft were then mounted on the Ariane duallaunch system (Sylda), with ECS-1 in the upper and Amsat in the lower position. The accompanying photographs show the final stages of integration. The launch-vehicle campaign started on 19 April. After the countdown rehearsal on 2 June, the ECS-1/Amsat/ Sylda composite was mated with the launcher on 7 June; the final launch countdown started on 15 June at 03.09 h local time (06 h.09 GMT). The Sylda The L6 flight allowed the first fully representative use of the Ariane duallaunch system Sylda (Systeme de Lancement Double Ariane). The system functioned as planned. The Sylda is an important element of the Ariane operational phase, since it allows the simultaneous launching of two spacecraft of mass up to 1200 kg into geosynchronous transfer orbit. Spacecraft of this size will constitute a major share of the world satellite market in the years to come, and the Sylda therefore increases Ariane's competitive position in this important sector of the launch market. Post-LS activities The successful L6 launch represents the culmination of a very significant effort initiated immediately after the L5 failure, as reported in the progress pages of ESA Bulletin No. 34. A Failure Enquiry Board was set up by ESA and CNES to establish the possible causes of the L5 third-stage malfunction and to recommend remedial action. The malfunction was rapidly traced to the either a breakdown in the turbopump gearing between the LH2 shaft and the LOX shaft, or a malfunctioning of the turbopump lubrication system. Measures have been taken to avoid the recurrence of this malfunction: for the gearing, by a tightening of the gear tolerances; and for the lubrication, by using a modified system on the Ariane third-stage engine. Considerable additional test time was accumulated with the turbopump (11 000 s) and with the third-stage engine (1300 s) prior to the L6 launch to validate the modifications. In parallel with this, all other launch elements were reviewed in depth to identify potential weak points. This exercise led to additional studies and tests, and to a strengthening of the test and acceptance procedures; it did not result in vehicle design changes. The Ariane L7 launch The next Ariane launch has been earmarked for Unit 7 of the Intelsat-V series, with lift-off scheduled for 15 September. The launch campaign starts on 8 July for the spacecraft, and on 4 August for the launcher.

ESA and Arianespace Sign Contracts for Launch of Four Satellites

Representatives of ESA and Arianespace have recently signed two separate contracts for the launch of four satellites. The first covers the launch in the second half of 1986 of the European Large Telecommunications Satellite (L-Sat), recently re-christened Olympus, on an Ariane-3 launcher. The second contract is a direct follow on to the recent signature of the EUMETSAT Convention and to the decision by participating countries to entrust ESA with the conduct of the Meteosat Operational Programme Signature of ESAIArianespace Contracts by (left) Mr. E. Mal/ett, ESA's Director of Application Programmes, and (right) Mr. Charles Bigot, Director General of Arianespace pending the entry into force of the Convention. This contract covers the launch of the three improved versions of Meteosat, in mid-1987, mid-1988 and at the end of 1990, respectively. It is planned to launch all three Meteosats on the Ariane-4 launcher now under development. The two contracts are worth over 130 million accounting units (or 1 thousand million French francs). They bring the Arianespace order book to 4.7 thousand million French francs, more than 40% of which relates to exports outside Europe. With these new orders, Ariane launcher production is already guaranteed up to launch L24. (31)

ECS - First Months in Orbit

Launch and early orbit operations ECS-1 was launched by Ariane (L6) on 16 June 1983, together with the radioamateur satellite Amsat. A brief description of the launch and the earlyorbit operations was given in the August 1983 issue of the Bulletin. The launcher provided an accurate transfer orbit and the satellite's apogee motor injected ECS1 into a nearly perfect drift orbit, the satellite arriving on station at 10"E longitude 21 days after launch, as planned. The operations during these 21 days followed the predetermined plan in copybook manner and were effectively a re-run of the many pre-Iaunch training simulations.

Giotto Ariane

In view of the Council decision to launch Giotto on an Ariane-1 vehicle, negotiations with Arianespace will be conducted during the remainder of 1983. Studies have already started, both internally and with Arianespace, to examine the feasibility and potential advantages of a direct-injection launch scenario. If there are indeed significant advantages, the current scenario for launching initially into a geostationary transfer orbit may be dropped in favour of direct injection into the comet transfer orbit. A final decision should be taken before the end of this year.

Ariane

ELA-2 The construction programme for the Second Ariane Launch Site (ELA-2) is proceeding on schedule, with a view to having the new launch complex available in October 1984. Work on the site at Kourou is proceeding as planned. The work force, which has numbered 450, is being reduced, now that the prinCipal civil-engineering tasks have been completed. In the launcher-preparation zone, infrastructure work on the Launch Centre and the air-conditioning plant is complete. Foundation work on the integration dock and the erection hall has been completed, and work has started on shielding and fitting out. In the launch zone, the umbilical tower is being fitted with an air-conditioning system and pipework. The main structure of the servicing gantry (~3000 t) has been erected up to the third stage. In the transfer zone, the double railway track has been laid and adjusted. The first mobile launch table in Ariane-3 configuration has been assembled and an initial preliminary test involving moving it along the track went smoothly. Construction of operational hardware is continuing in Europe. Factory acceptance of the cryogenic arms and testing of the electrical command and control systems have begun. There have been some performance problems with the fluids command and control system, which has been undergoing factory acceptance since July, and these will need to be solved before shipment to Guiana. The release system is a critical item in the programme installation of this system, which is currently being set up, is still planned for the summer of 1984.

Launch contract signed for Marecs-B2

On the basis of a contract signed by representatives of the European Space Agency and Arianespace at the end of July, the Agency's Marecs-B2 satellite will be put into orbit by one of the first Ariane-3 launches in the first half of 1984. Marecs-B2, with a launch mass of 1050 kg, will replace an earlier version of the same spacecraft which was lost following a launch failure on 10 September 1982. The Ariane-3 vehicle is designed to place a single satellite of the 2500 kg class, or two of the 1200 kg 62 class in a dual launch configuration, into geosynchronous transfer orbit. Marecs-B2 will be stationed at 177HE above the Pacific Ocean. As part of the Inmarsat global maritime communication system, the Marecs satellites provide shipping and offshore industries with access to international public telephone (more than 40 channels) and telex networks, and with facsimile and data-transmission facilities. They also provide a unique facility for handling priority messages for maritime distress and safety services.

Launch contract signed for ECS-3

On 29 September 1983, ESA's Director of Applications Programmes, Mr. E. Mallet!, and the Director General of Arianespace, Mr. C. Bigot, signed a contract for the launch of the third European Communications Satellite, ECS-3. This launch, scheduled for August 1985 on an Ariane-3 vehicle, will follow that of ECS-2 in May 1984, for which the launch contract has already been signed. The first ECS satellite was successfully launched by Ariane L6 on 16 June 1983 and is expected to become operational shortly as an integral part of the Eutelsat system (see elsewhere in this Bulletin). This new contract brings the Arianespace order book up to a total of 5.2 thousand million French francs, for the launch of 25 spacecraft. In the telecommunications field alone, the European Space Agency has now signed contracts worth 1.6 thousand million French francs for the launch of two Marecs satellites for maritime communications, three ECS satellites, and Olympus, the European large telecommunication and direct TV satellite.

Ariane Launches

Intelsat-V Intelsat-V F7 was launched successfully, as Ariane's first purely commercial payload, from ESA's Kourou launch base in French Guiana at 21.45 local time on 18 October (45 minutes after midnight on 19 October GMT). The Ariane launch vehicle injected the International Telecommunications Satellite Organisation's two-ton satellite into an orbit with a perigee of 183 km (against 185.4 km specification) and an apogee of 36 158 km (agai nst 35987 km specification) Geostationary Transfer Orbit (GTO) was achieved at 2.30 local time. The launch of Intelsat-V F7, which was originally scheduled for 15 September, had been postponed at the customer's request to allow them to investigate a potential problem in the satellite's maritime communications subsystem. The launch of Intelsat-V F8, also aboard a European Ariane vehicle, remains on schedule for mid-December. (32)

Ariane

Ariane-2/3 Follow-On Development The Ariane-2/3 Follow-On Development programme is in its final phase. All the system studies and tests have been completed. A final computer run will be made to verify the values of the general loads.

Qualification of the various parts of the launcher is nearing completion. The meetings of the qualification boards started in the last quarter of 1983 and will end after the last test of the third-stage propulsion system scheduled for midFebruary 1984. As the programme does not provide for a qualification flight, qualification will be certified on the basis of the results obtained from the ground tests. The last meeting of the Ground Qualification Board is planned for March 1984. In the course of November 1983, a CSG facilities validation exercise was carried out in order to test their ability to attach the boosters to the structure of the first stage. The results obtained were very satisfactory. The Flight Readiness Review (FRR) for the first Ariane-3 launcher has been scheduled so that the first launch in Ariane-3 configuration from ELA-1 can take place after the end of May 1984. The first three launchers, intended for operational purposes, will however carry technological telemetry to provide additional data on the functioning of the various systems and equipment. For the first time, the chamber pressure setting for the Viking engines (first and second stages) will be raised to 58.5 bar (instead of the 53.5 bar for Ariane-1) which, with the assistance of the solidpropellant boosters, will allow a payload of 250 kg to be placed in transfer orbit. (33)

Ariane Launches Second Intelsat-V

The International Telecommunications Satellite Organisation's Intelsat-V F8 was successfully launched by Ariane vff, from ESA's Kourou launch base in French Guiana at 21.50 local time on 4 March (00.50 on 5 March GMT).

Separation of the satellite was nominal and, according to first reports received from Intelsat, it was placed by Ariane in a near-perfect transfer orbit, with a perigee of 186.5 km (against 184.9 specification), an apogee of 36045 km (against 35988 specification) and an inclination of 8.530 (against 8.500 specification). Satellite orbital operations proceeded nominally, and the apogee boost motor was fired at 11.10 GMT on 7 March.This was the second successful launch of an Intelsat-V satellite by Ariane under the auspices of ESA. The launch of the third Intelsat-V, the last in ESA's 'promotion series', has been subcontracted to Arianespace, who then take over all future launch-activity responsibility. The accompanying table shows the provisional schedule for future Ariane launches. (34)

1979-1983: the Ariane era

By 1979, ESA had an ambitious launch programme of its own based on the imminent availability of its European Ariane launcher; the schedule included: Meteosat-2, the second European weather satellite, to be injected in dual-launch configuration with ISRO's Apple satellite Marecs-A and B, the first European maritime communications satellites Sirio-2, a combined scientific and meteorological data-relay satellite ECS-1, the first operational European communications satellite (OTS was designated a test satellite) Exosat, the first European X-ray astronomical satellite. To support all of the above missions (except Meteosat-2), it was necessary to construct and install the dedicated operational facilities outlined in Table 1. All of these facilities were installed, tested and operationally validated between 1978 and 1983, prior to the launch of each satellite. In 1983, because of launch delays, the launch and early orbit phases of two satellites ~ad to be supported within three weeks of each other and both operated from the ESOC-OCC at Darmstadt - Exosat being launched from Western Test Range, California on 26 May and ECS-1 from Kourou on 16 June (Fig. 6). On 12 October 1983, when ECS-1 was officially handed over to service with Eutelsat, ESA was operating a total of nine satellites: Meteosat-1 and 2 (meteorological), OTS-2, Marecs-A, ECS-1 (communications), and ISEE-B, Geos-2, IUE and Exosat (scientific) - the majority being geostationary satellites requiring dedicated ground stations and control rooms, and 24 h per day surveillance The contrast between the early days of the Estrack network, with 50 baud telex communications and the present ESA worldwide network equipped with 9.6 kbitls digital data links, providing realtime command and telemetry functions, is certainly extraordinary in its degree.

Ariane

Ariane-3 launcher ground qualified Since the Ariane-3 development programme does not include any test flights, qualification consisted of carrYing out two types of check on the ground: programmes & operations studies, calculations and tests on the elements that are new or modified compared to Ariane-1 (solid boosters, thrust increase for the first- and second-stage Viking engines, increased burn-time for the third stage, strengthening of the structures and biconic fairing); checking the capability of the unmodified elements to fulfil the mission under Ariane-3 conditions (duration and environment). When about 90% of the qualification reviews of the various elements had taken place, ESA and CNES set up a Ground Qualification ReView Board (GQRB) consisting of members of ESA, CNES and Arianespace, plus independent experts whose task was to assess the status of the launcher qualification. The GQRB, whose work started in September 1983, submitted a final report on 12 June 1984 to the ESA and CNES General Directorates, which approved it and presented it to the Ariane Launcher Programme Board on 14 June 1984. The latter declared the launcher qualified subject to the reservations made in the GQRB report, and in particular to the need for some of them to be lifted before the first Ariane-3 launch scheduled for 4 August 1984. (35)

Giotto On-Course for Halley's Comet

On 2 July 1985, the Ariane-1 launch vehicle V-14 carried the Giotto spacecraft into orbit after some 15 minutes of powered flight. Lift-off from Kourou, French Guiana, occurred at 11 :23:16 GMT. Twenty-two minutes later the first telemetry signals from Giotto were received at the Estrack station at Malindi in Kenya. From this point in the mission until the third perigee, some 32 h later, telemetry data acquired from Malindi and two other Estrack stations, at Carnarvon, Western Australia and Kourou, French Guiana, flowed via digital data links (for processing by the ESOC computers and display of engineering data to the Mission Control Team) to the ESOC Operations Control Centre at Darmstadt, West Germany, as Giotto orbited the Earth in Geostationary Transfer Orbit (GTO). This period of 32 h was an extremely busy one, for the members of the ESOC Mission Control Team, the Giotto Project's spacecraft support team, and the staff of the tracking stations. It was also a stressful period because, for the first time, the newly designed and tested OCC was being used for active mission control together with newly installed S-band tracking stations plus a newly installed data-packet-switched communications network (see pages 24-31 of this issue). In the event, the new systems performed well and all flight events were carried out according to the Flight Operations Plan (FOP). In order to place Giotto into the desired heliocentric interplanetary trajectory, it was necessary to determine the orbital parameters of the GTO and the attitude of the Giotto spacecraft precisely and to increase the spacecraft's spin rate prior to igniting its solid-propellant Transfer Propulsion System (TPS). Between approximately 1200 GMT on 2 July and 19:23:47 GMT on 3 July, the following activities and manoeuvres were carried out using the telemetry, tracking and telecommand facilities of the Estrack S-band network: determination of GTO orbital parameters determination of spacecraft attitude three slew manoeuvres to place the spacecraft into the optimum attitude for firing the TPS; spin-up to 90 rpm, in two stages, in order to stabilise the spacecraft prior to TPS ignition. The TPS firing took place precisely as planned by means of time-tagged stored commands, while Giotto was out of contact of the network during third perigee passage, and then confirmed by telemetry via the Kourou tracking station. Within three hours, orbit determination revealed that Giotto was now speeding at some 12 km/s away from the Earth, exactly on course for its rendezvous with Halley's Comet. The orbit-insertion manoeuvre was so precise that the need for any orbital adjustment using the onboard thrusters was obviated. During the next four days, the Giotto spacecraft systems were checked out and configured for the long journey (700 000 000 km) to the Comet. The spacecraft's spin rate was adjusted to 15 rpm, and on 6 July 1985 at 10.47 GMT its high-gain antenna was mechanically released and despun. The telemetry signal received at the tracking stations increased in strength by a factor of more than 1000 as this antenna was activated. By 20.00 GMT on the same day, the testing of all Giotto platform systems was complete and the start of the mission's long 'cruisephase' to the Comet had begun. As of 15 July 1985, the Giotto spacecraft is 3427000 km from Earth and moving away from us at 13000 km/h. On 15 July, a radio command transmission to Giotto required 11 .5 s to reach the spacecraft from Earth. On 13 March 1986 a command sent to Giotto will travel at the speed of light for 500 s before arriving at the spacecraft. Europe's first interplanetary scientific probe is now on course for rendezvous with the Comet at 24.00 GMT on 13 March 1986. (36)

Ariane Launch Calendar

The Ariane V16 launch, using the last Ariane-1 launcher, will place into heliosynchronous orbit the French Earth observation satellite SPOT and the Swedish scientific satellite Viking. The V16 launch, previously postponed until 16 January 1986 due to the failure of V15, has now been rescheduled for 21 February following the discovery of a crack in the 2nd-stage water tank. While experts investigate the cause of this, the defective tank will be replaced by a new one. This operation will take place on 23 January. The updated timetable for V16 and subsequent launches is currently as follows: V16: SPOT & Viking (Ariane-1 from pad ELA-1) • 21 February V17: G-Star 2 & Brasilsat (Ariane-3 from pad ELA-2) • 12 March V18: Intelsat-V F14 (Ariane-2 from pad ELA-1) • Approx. 2.5 months after V16, i.e. end April V19: ECS-4 & Spacenet 4 (Ariane-3 from ELA-1) • Two months after V18, i.e. approx. end June V20: TV-Sat (Ariane-2 from ELA-1) • Two months after V19, i.e. approx. end August (Arianespace is studying the possibility of launching V20 from ELA-2). (37)

Birth of the space propulsion industry

At that time the company’s main space activities concerned small solid propulsion, with the apogee kick motor for the Sirio 1 test satellite and the separation rockets for the stages of Ariane 1, the new European launch vehicle.

An early assignment was a project for the newly created European Space Agency: to develop software to manage the solid propellant’s internal ballistics for the MAGE apogee kick motor developed for future geostationary satellites. As a design engineer on the MAGE programme, he later travelled to the sites of a nascent European solid propulsion industry: Bordeaux, Munich and ESTEC, ESA’s technical centre in Noordwijk, the Netherlands.

The first MAGE motor flew on the third Ariane 1 launch that on 19 June 1981 boosted the Meteosat 2 satellite into geostationary orbit. Some 18 motors were flown in 1997 on Ariane launchers and the Space Shuttle. One is lost in a launch failure and another cannot be ignited due to an electrical problem unrelated to the motor itself. The remaining 16 do their job perfectly, putting Europe’s first communication and meteorology satellites into operational orbit and sending the Giotto probe on its way towards a dramatic encounter with the nucleus of Comet Halley comet in 1986. (38)