This article describes the circumstances that led Europe to go ahead with the Ariane Program in 1973 and subsequently to decide on a series of follow-up versions, up to and including Ariane 5. The article also describes the main management principles adopted that have varied little in 20 years and includes brief details of the successive launcher configurations from Ariane 1 to Ariane 5.
Page of History
In 1972, the European space community was in a state of crisis. It is interesting to look back at the situation which preceded this crisis to understand the reasons that led to the decision to proceed with the Ariane program and have subsequently guided the program up to the present time.
The first European initiative in the launcher field was taken by the United Kingdom in 1961, which put forward proposals to France, and later Germany, for manufacturing a launcher based on the Blue Streak, roughly similar to the American Atlas, designed and developed up to that time as a first missile stage. The Blue Streak stage was available, following a change of British policy in this domain. This was the basis for the Europa 1 project that was designed to place payloads of 1 metric ton into low orbit.
Unfortunately, the organization set up already contained the seeds of failure because the United Kingdom was supplying a first stage, whose characteristics were frozen, whereas France took on responsibility for the second stage, and Germany for the third stage. Each adopted its own technology and retained an autonomous stage, creating a total absence of anything resembling a system study or an attempt at general optimization. Furthermore, the program was created by cooperation among sovereign states, coordinated by a secretariat under the direction of a diplomat.
By 1966, there was no lack of technical and financial problems. The British threatened to withdraw from the project, which they did two years later. Therefore, the program was obliged to buy the first stage directly from United Kingdom industrial firms. The program was redirected toward geostationary orbit launches (a satellite placed into geostationary orbit appears to be in a fixed position with respect to Earth), and a solid propellant fourth stage was added. At the same time, the Australian launch range was abandoned in favor of a new launch center to be constructed in French Guiana. The payload specification was then 270 kg in geostationary orbit, and the new program was dubbed Europa 2.
At the same time, geostationary orbit application prospects became clearer, and in 1970, Europe decided to commence studies and predevelopment work on a more powerful launcher (Europa 3) that could place 1500 kg payloads into geostationary transfer orbit (GTO). Unfortunately, the lessons to be learned from the Europa 1 failures were ignored insofar as program organization was concerned. Furthermore, although the first-stage technical configuration remained prudent, based on experience acquired by France from its Diamant program (first unit acquired in November 1965), the second stage was much too ambitious for Europe and involved a high-pressure liquid hydrogen/liquid oxygen topping cycle engine delivering 200 kN of thrust.
The period from 1969 to July 1973 was difficult for the European space community for the following reasons:
• In the technical context, Europa 1 launches F7, F8, and F9, made from the Australian base, all failed. The launcher on flight F11, the first launch Europa 2 made from the new range in French Guiana, exploded 150 seconds after lift-off on 5 November 1971.
• In the political sphere, Germany decided to withdraw from the Europa 2 program in December 1972 (which led to stopping that program in April 1973) and temporarily suspend its participation in the Europa 3 program. The Germans were also considerably attracted by collaborative proposals made by a NASA glowing from the success of the Apollo program. Germany considered that European efforts in the space transportation sector could be limited essentially to technological development, in parallel with major participation in the post-Apollo program. Nevertheless, negotiations between NASA and European representatives on that subject were difficult and engendered frustrations in Europe. At first invited to be involved in developing specific Shuttle hardware, the European contribution progressively narrowed to a science module that would fit into the Shuttle cargo bay.
France adopted a different approach that expressed a triple objective:
• acquisition of absolute control of space applications;
• founding of this control on an autonomous launch capability, with particular reference to geostationary satellites; and
• adoption of these first two objectives by its European partners to assemble sufficient financial capacity and sufficient volume to make production feasible.
However, these French ambitions were momentarily weakened by two launch failures in the Diamant national program in 1971 and 1973, despite the fact that they followed a run of six successful launches.
Finally, the Ariane program may well owe its very existence to the difficult negotiations undertaken with NASA in connection with the launch of the two “Symphonie” experimental telecommunications satellites. The extremely harsh conditions imposed on the German and French negotiators, including an embargo on using these two satellites for any commercial purpose in particular strengthened the determination to achieve the autonomy proposed by France. European agreement was finally achieved in July 1973, following a period of intensive negotiation.
It was decided to embark on three simultaneous programs:
* the “L3S” heavy launch program proposed by France at the end of 1972 (this program was renamed “Ariane” shortly afterward);
* the Spacelab program in cooperation with NASA, backed by Germany; and
* the “Marecs” maritime telecommunications program backed by the United Kingdom.
The principles of setting up the European Space Agency were also defined, and the Agency was officially formed in 1975.
The commitment made by France in connection with this agreement was very considerable and corresponded to more than 60% funding for the program, plus an undertaking to fund excess cost above 120% of the initial figure of MFF 2,060 (about $ 447 M 1973) up to a maximum of 15% of this figure. In exchange, France obtained agreement that the European Space Agency (ESA) would delegate management of the program to the French Space Agency (CNES). The obligation to ensure a workload return for each participating country in proportion to its contribution was and is indeed a particular aspect ofthe ESA programs to be emphasized (Fig. 1).
The Early Days of Ariane
The performance objective for the L3S launcher was rather ambitious, based on the conviction that telecommunication satellites mass values would continue to increase and that Europe had to prepare itself for this evolution. Although the view was unanimous that the geostationary orbit was the most promising for application satellites, the figure of 1500 kg had already been the subject of arguments which were to be reopened at regular intervals when performance enhancements were proposed. The conflict was (and still is) between those who predicted a reduction in satellite mass values under the combined effect of electronic circuit miniaturization and the enhanced performance characteristics of satellite onboard propulsion systems and those who predicted a continued increase in mass values under the effect of traffic growth, congestion of the geostationary orbit, and the resultant reduction in in-orbit transponder cost. Finally, the payload mass objective adopted for the Europa 3 launcher was confirmed. It was set at 800 kg for geostationary orbits, corresponding to about 1500 kg in geostationary transfer orbit (about 200 km by 36,000 km). This represented almost twice the performance of the American Delta launcher, which had previously launched most application satellites for the Western world. This performance was comparable with that of the Atlas-Centaur, which was used at the time only for few heavy satellites. This decision provided a substantial growth potential for European satellites and demonstrated a determination to design the Ariane launcher for an extended lifetime.
| Contribution % | ||
| Initial | Final | |
| 63.87 | 73.55 | |
| Germany | 20^2 | 11.47 |
| Belgium | 5.00 | 4.71 |
| United Kingdom | 2.47 | 2.27 |
| Spain | 2.00 | 2.38 |
| Netherlands | 2.00 | 1.64 |
| Italy | 1.74 | 1.58 |
| Switzerland | 1.2 | 0.83 |
| Sweden | 1.1 | 1.05 |
| Denmark | 0.5 | 0.52 |
Figure 1. National financial contributions. Financial management of the Ariane program had to ensure a fair return. This meant that each country was to receive a volume of business in proportion to its financial contribution. It was also expected that these figures would remain at the same level for the subsequent production phase. Already difficult to apply because it concerns an objective which is both final—including contingencies—and subject to examination when budgets are voted on for the following year, this constraint was further complicated for Ariane 1 by updating rules which varied from one country to another. For example, Belgium, France, and Switzerland made commitments based on a percentage of initial development cost, whereas Germany undertook to make an annual contribution, expressed in its national currency, which would be revised only once in midprogram, according to monetary parities and observed inflation rates. The United Kingdom was involved only through a specific agreement with France. Furthermore, inflation rates at the time were in double digits, and application of the rule of fair return, while fully understandable in principle, was far from easy. The figure compares the financial contributions at the start of the program with the final contributions.
Past experience, as mentioned at the beginning of this article, demonstrated the absolute need for imposing three major principles at the very start of the program, namely, a genuine system approach, strong management, and a simple design.
The System Approach. This made it necessary to regard the Ariane launcher (and its ground support facilities) as a single entity, not merely a stack of independently designed stages; to base the development of these stages and other subsystems on the results of studies conducted at the highest level (trajectory, flight mechanics, general loads, thermal, guidance and control, etc.); to design electrical systems in terms of the complete launcher, installing only actuator devices and equipment specific to its flight phase in each stage; and to impose design rules common to each discipline. For example, it was in this way that the need for POGO control systems was demonstrated at the beginning of the development phase. This meant that corrector devices could be designed and integrated into the propulsion systems at the outset. This was of even greater interest because Ariane was the first launcher designed entirely for geostationary orbit missions, in contrast to the other launchers existing at the time that derived to a greater or lesser degree from ballistic missiles. CNES entrusted a specific contractor (Aerospatiale) with this task and also associated Aerospatiale with the reviews conducted during the development of the various launcher systems and subsystems.
Strong Management. After obtaining delegated management authority, CNES spent the first year of the program establishing the basis for a strong management structure founded on a number of specific principles:
* unique management link between CNES, main contractors and other contractors;
* clearly defined industrial organization, with precise definition of the tasks allocated to each party. For the Ariane 1 program, CNES used the same French level 1 contractors as for its Diamant program, thus ensuring design unity in the main disciplines. These were Aerospatiale for the launcher stages, SEP (Societe Europeenne de Propulsion) for the propulsion systems, Matra for the vehicle equipment bay, and Air Liquide for the cryogenic third-stage tanks. Other contractors subsequently achieved level 1 status, including Contraves (Switzerland) for the fairing, DASA (Germany) for the second-stage and then the Ariane 4 liquid propellant boosters, Fiat-BPD (Italy) for the Ariane 3 and 4 boosters, etc.;
* a common understanding of the content of work to be done by each contractor, including achievements expected and reports to be submitted.
These principles led CNES to issue a set of management specifications applicable to each launcher system, subsystem and component, and also its ground facilities. These specifications covered overall planning and milestones, work breakdown structure, industrial organization, technical work coordination, schedule and cost reporting, monitoring of the development of critical elements, quality, and reliability. Application of such specifications in the context of a program involving 10 countries that differ in language and culture was a new departure for Europe and generated initial difficulties with companies each of which had their own particular methods of working. Nevertheless, this proved essential to maintain both visibility and coherence. These basic principles were also applied to the Ariane complementary development programs. They proved their effectiveness and made it possible to achieve effective control of technical development, costs, and time schedules.
Simple Design. This final point meant that the Ariane design should be based only on technologies which were either already available or involved only low development risks. The Europa 3 first stage had been designed on this principle, applying technologies and experience acquired with the first stage of the French Diamant launcher. This stage was powered by four rugged “Viking” storable propellant engines, for which the prototype, designed in France, had been tested with very encouraging results before this program was terminated. Therefore, the Europa 3 stage was selected as the basis for the first stage of the Ariane launcher.
This could not be the case for the Europa 3 second stage that was considered much too ambitious for Europe in 1973. However, although a number of difficulties were to be anticipated if liquid hydrogen and liquid oxygen were used, the efficiency of these propellants would make it possible to reduce the lift-off mass of the launcher by a factor of almost 2. Furthermore, a certain amount of experience had been acquired on a low-thrust gas generator cycle prototype engine developed within the French national program. Work conducted in France and Germany during the Europa 3 program meant that this approach could be adopted with a satisfactory level of confidence, provided that there was no major deviation from the experience acquired from the engine and propellant and fluid system. Thus, the tanks were designed for 8 tons of propellant and a maximum possible diameter using available tooling not to exceed 2.6 m. Having made this choice, it was found necessary to add a second, intermediate stage to achieve the performance target. This stage was designed on the basis of elements and technologies developed for the first or third stage, including the Viking engine used in the first stage (with a suitably adapted nozzle) and the light alloy tanks of the third stage.
Ariane 1 Launcher
Ariane 1 was a three-stage launcher that had a total height of 47.8 m and a lift-off mass of 210 tons. The first stage had a dry mass of 13.3 tons, a height of 18.4 m and a diameter of 3.8 m. It had four Viking 5 engines (Fig. 2) that developed 2500 kN thrust on lift-off. Burn-time in flight was 146 s. The 148 tons of propel-lant (UDMH and N2O4) were contained in two identical steel tanks protected again internal corrosion by an aluminum layer and connected via a cylindrical skirt. The four turbopump Viking engines were mounted symmetrically on a thrust frame and articulated in pairs on two orthogonal axes to provide for three-axis attitude control. An annular water tank, located inside the propulsion bay, provided for cooling the gas obtained from the engine gas generators that was used to pressurize the propellant tanks and supply the hydraulic motors of the attitude control actuator systems. Four fins with a surface area of 2 m2 provided aerodynamic stability. This first stage was designed to destruct approximately 30 seconds after stage one to two separation.
Figure 2. Viking engine flow diagram. The Viking engine uses the gas generator cycle, and the gas generator itself operates at a stoichiometric ratio. The gas produced is cooled by injecting water to reduce gas temperature to values compatible with the turbopump turbine, pressurization of the main propellant tanks, and operation of the power pack for the hydraulic servoactuators. The three pumps (UDMH, N2O4, and water) are mounted on a single shaft that rotates at 10,000 rpm. (For Ariane 1, the chamber pressure was limited to 53.5 bars.) The hydropneumatic regulation system slaves the combustion pressure to a reference pressure value by adjusting the gas generator feed and, thus, the rotational speed of the turbopump. The mixture ratio is maintained at a constant level by a regulator that equalizes propellant pressures before injection into the combustion chamber. Engine ignition is induced by the pressure in the propellant tanks. When the valves open, the propellants, which are hypergolic, are delivered to the combustion chamber and gas generator and ignite spontaneously. The turbopump speed then builds up to the value set by the regulating system in 1.3 s.
The second stage had a dry mass of 3.13 tons (excluding the interstage conical skirt and the jettisonable acceleration rockets), its height was 11.6 m, and its diameter was 2.6 m. It had a single Viking 4 engine that developed 740 kN thrust in vacuum for a burn-time of 136 s. The motor was linked to the conical thrust frame via a gimbal with two degrees of freedom, that provided pitch and yaw control. Auxiliary nozzles supplied with hot gas from the engine gas generator provided the roll control function. The two aluminum alloy tanks that had a common intermediate bulkhead were pressurized with helium gas (3.5 bars) and contained 34.1 tons of propellant (UDMH and N2O4). The second stage was also designed to destruct about 30 seconds after stage two to three separation. During the prelaunch waiting period on the pad, a thermal shroud, ventilated with cold air, which restricted heat exchange between the propellants and the environment, protected the second-stage tanks. This shroud was jettisoned on launcher lift-off.
The third stage weighed 1.164 tons dry, was 9.08 m high, and had a diameter of 2.6 m. This was the first cryogenic stage produced in Europe. It was equipped with a type HM7A engine that developed 62 kN thrust in vacuum for a burn-time of 545 s. This engine was designed by Societe Europeenne de Propulsion (SEP), based on experience acquired with an earlier cryogenic engine, the HM 4, which delivered 40 kN thrust, and tested over the period 1962-1969. The HM 7a engine uses the conventional gas generator cycle technology, and achieves a specific impulse of 443 s with a mixture ratio of 4.43 at pump inlet. The combustion chamber is supplied with propellants pressurized by a turbopump, via a set of injection valves. The pump turbine is driven by gas supplied by a generator, the latter receiving a small proportion of propellant tapped off at the pump outlet. The liquid hydrogen and liquid oxygen tanks that contained a total of8.23 tons of propellant were made of an aluminum alloy and had a common intermediate bulkhead (double skin under vacuum). The tanks were covered with external thermal protection to avoid warming the propellant. Both tanks were pressurized in flight, using hydrogen gas tapped at the outlet from the regenerative chamber and helium. The motor was linked to the conical thrust frame via a gimbal that provided pitch and yaw control. Auxiliary nozzles ejecting hydrogen gas were used for roll control.
The stages were separated by pyrotechnic cutter devices fitted on the rear skirts of the second and third stages. The separating stages were distanced from each other by retrorockets incorporated in the lower stage and acceleration rockets mounted on the upper stage. Stage 1 to stage 2 separation was controlled by the onboard computer, on detection of first-stage thrust decay (propellant exhaustion). Stage 2 to stage 3 separation was controlled by the onboard computer when the second-stage speed increase reached 1500 m/s.
The vehicle equipment bay (VEB) weighed 316 kg, had a diameter of 2.6 m, and was of 1.15 m high. Mounted on the third stage, the VEB contained the electronic equipment of the launcher and also served as a base for the payload and fairing. In addition to the onboard computer, the VEB housed all of the electrical equipment required for executing the launcher mission, namely, the sequencing unit, guidance and navigation control, and the location, safety, and telemetry systems. Only the power systems and actuator devices were located elsewhere in the launcher stages.
The two half-fairings were ejected parallel to the main axis of the launcher under the control of the onboard computer, as soon as the calculated thermal flux dropped below the specified level. The fairing was jettisoned by two pyrotechnic systems, a horizontal system at the interface with the VEB and a vertical system which also served to impart horizontal velocity to each half-fairing. The fairing was 3 m in diameter and was compatible with Atlas-Centaur class satellites.
The launch facilities in French Guiana (ELA 1:Ariane launch site No1) were designed to make use of the earlier investment for the Europa 2 launcher. The stages were erected and assembled directly on the launch pad. A mobile gantry sheltered the launcher and provided for assembly of the payload with the launcher and closure of the fairing under clean room conditions. The gantry withdrawal commenced 6 hours before lift-off. The main fueling of the third stage with liquid hydrogen and liquid oxygen and final topping off were performed using a cryogenic arm system carrying a set of umbilical valve plates. Disconnection and retraction of the arms commenced at time T — 3 s. Control of the launcher on the pad was fully automatic from time T — 6 minutes. The launcher was then controlled by two ground computers, one for the electrical systems and the other for the propellant and fluid systems. The two computers also crosschecked each other. The first stage engine was ignited at time T, and the lift-off command signal was sent at T + 3 s, following satisfactory verification of the Viking engines. The first flight took place on Christmas eve 1979 (Fig. 3).
When the Ariane program was first initiated in 1973, few imagined a commercial career for the launcher. Things began to move in 1976 and led to a series of actions relating to launcher performance, production, and marketing. As regards performance, the idea was to propose Ariane for launching the Intelsat satellites. The success of such a venture would represent both an exceptional reference for the program and strong motivation for the players involved. However, the first task was to augment performance objectives up to 1600 kg minimum in GTO. Fortunately the prudent approach adopted for the basic definition of the launcher, combined with the results of the first flight, demonstrated propulsion performance in excess of specifications and made it possible to exceed the initial objective and achieve a figure of 1850 kg.
As far as production was concerned, it was obvious that there was no hope of selling the launcher if one waited to receive orders before commencing manufacture. It was on these lines that discussions were opened at the European Space Agency, and a promotional phase was duly adopted. Apart from the production of six launchers, the objective of this phase was to achieve full operational qualification, develop and validate the dual launch capability, and adapt the launcher comprehensively to meet user needs by providing for the construction of payload preparation facilities in French Guiana. Analysis of marketing aspects led to two actions: initiation of the Ariane 3 program and formation of Arianespace.
Figure 3. Launch L0 1 (24 December 1979). The decision to proceed directly to flight tests, using three active stages in the final flight configuration, was made following very lengthy discussions at the start of the Ariane program. This was in total opposition to the highly progressive, but also extremely costly approach adopted for the Europa 2 program. Initial ignition occurred on 15 December 1979, following a faultless countdown. Unfortunately, the ground computers did not authorize liftoff, and the engines shut down automatically at time T + 10 s and aborted the launch. Subsequent analysis showed that an explosion in a small measurement pipe had damaged the sensors, whose signals were used for operational diagnosis of the engines. The case of an aborted launch, for which the probability was infinitely small, had, nevertheless, been taken into account during the development phase. Procedures for return to flight configuration had been written and validated by tests conducted in Europe. This allowed restarting the count on 23 December, following 8 days of round-the-clock work. Technical problems, combined with adverse meteorological conditions, prevented a launch on 23 December. On the following day, Ariane made a practically faultless launch. Only two anomalies were identified, and these were corrected before the following flight. These concerned minor pollution of the payload, caused by the second-stage retrorockets and low amplitude vibrations (POGO effect) at the end of the second-stage flight. The first Ariane launch thus took place only 6 months after the initial target date set in 1973.
Ariane 3 Program
At the end of the 1970s, the application satellite market was organized around two launcher groups. These were the Delta launchers, for which GTO performance in 1978 was around 1200 kg and the Atlas-Centaur launchers whose performance figure was close to 1800 kg. Apart from the Intelsat satellites, all other payloads corresponded to the Delta class, for which Ariane could not pretend to be competitive in its actual state. The idea then emerged to adapt the launcher to enable it to launch two Delta class satellites simultaneously, thus raising the performance objective to twice 1200 kg, plus the mass of the dual launch structure designed to isolate the two satellites from each other. This corresponded to a total GTO mass of 2500 kg. The launcher also had to be a competitive with the Atlas-Centaur.
This program was proposed by France to the European Space Agency and was approved in July 1980, despite the failure of the second Ariane flight (L02) in May of that year. CNES was charged with managing the Ariane 3 program, under conditions very similar to those for Ariane 1.
Analysis conducted from 1976 and aimed at increasing the performance of Ariane 1, had in fact identified modifications whose feasibility was checked out during the final development test phase. These modifications were adopted in part or in full for the Ariane 3 program, in accordance with the policy of minimizing development risks. The addition of two solid propellant boosters achieved the objective of competitiveness, with Ariane 3 for two satellites of 1200 kg each, and Ariane 2, with no boosters, for satellites of 2000 kg.
The second launch in May 1980 resulted in a failure. Destruction of the launcher occurred at T + 63.75 s, as a result of high frequency combustion instability in one of the first-stage Viking engines, 2.75 seconds after liftoff. Corrective measures essentially comprised modifying the propellant injection orifices in the combustion chamber. A total of 95 tests that represented 4300 seconds of burn-time were conducted to qualify the new injector under conditions substantially more severe than those encountered in actual flight. The Ariane 1 development phase terminated in December 1981, following three successes out of the four launches made. The total cost of the program was within 120% of the initial estimate.
These performance objectives were achieved on the basis of the Ariane 1 launcher by introducing the following modifications:
* increased thrust for the first- and second-stage Viking engines by augmenting combustion pressure by 10% (53.5 bar to 58.5 bar). This was obtained mainly by adding hydrazine hydrate to UDMH;
* adding two solid propellant boosters with a unit thrust of 600 kN and a burn-time of about 40 seconds. The low level of performance sensitivity to the structural mass of the boosters made it possible to adopt extremely prudent technical solutions;
* increase in the third-stage propellant load from 8 to 10 tons;
* enhancement of Stage 3 performance by increasing combustion chamber pressure by 5 bar and stretching the nozzle by 200 mm (HM7 B);
* adaptation of the SYLDA dual-launch system and the fairing to the volume required for 1200 kg-class-satellites.
Figure 4 shows a SYLDA dual-launch structure and half a fairing. The lower satellite is placed inside an egg-shaped compartment. Protected by the fairing, this carbon fiber structure is subject to reduced loads by comparison with an external structure. Furthermore, the operational constraints induced by the need to carry two satellites remain within acceptable limits.
The SYLDA structure makes it possible to achieve complete separation of the two payloads. The long orientation sequence, spin-up phase and separation of the satellites and upper part of the SYLDA, are achieved by the SCAR attitude and roll control system, using third stage pressurization hydrogen gas to operate this system.
Tested successfully on the sixth launch, this concept was unquestionably one of the keys to the success of Ariane. The cost of a launcher is not proportional to its size, and a number of functions must exist whether the launcher is small or large. A dual-launch capability, taking advantage of this scale effect, represents a major competitive plus factor.
The industrial organization was the same as that for the Ariane 1 program; the only main modification was an increase in the Italian contribution, which led to entrusting the development of the solid propellant boosters to FIAT-BPD. The first Ariane 3 was launched successfully in mid-1984.
Arianespace
The lengthy discussions which took place at the European Space Agency and led to the decision to go ahead with production of six Ariane 1 launchers, also demonstrated that this multinational organization, whose basic purpose was research and development, was not suitable for engaging in commercial and production activities. At the same time, the promoters of Ariane reached another paramount conclusion. For European autonomy to be effective in access to the geostationary orbit, the Ariane launcher had to be credible. To this end, it had to be both reliable and available, in other words produced in sufficient quantities and in advance of actual launch needs. This called for world-scale marketing, and at the same time made it possible to spread the fixed production costs over a larger production volume, and thus reduce the cost of the autonomy-related strategy (in contrast to its American competitors, Ariane did not have a major captive market for launching governmental satellites, to which all or part of the fixed costs could be allocated). Only a private commercial entity, responsible for both Ariane production, marketing, and launch operations, could take up this challenge.
Figure 4. SYLDA, Ariane dual-launch system.
In December 1977, Frederic d’Allest, then head of the CNES launcher division and future Chairman and CEO of Arianespace, proposed forming a company to market the Ariane launcher. Arianespace was incorporated on 26 March 1980, well before qualification of the Ariane 1 program. Arianespace is a French business corporation. Its capital is held by CNES, the leading European manufacturers that participate in Ariane production, and a number of banks. Ari-anespace is responsible for producing, launching, and marketing the Ariane launcher or launchers, whose development and qualification have been, are, or will be conducted by the European Space Agency.
The first Arianespace commercial launch was made successfully on 22 May 1984, inaugurating the very first commercial space transportation service.
Quality
A number of technical difficulties were encountered during the early days of the Ariane program, resulting from residual design problems or omissions in the production files. The European launcher industry was in its infancy was in the process of discovering the problems inherent in launcher batch production, and had to acquire expertise in this demanding discipline. Major efforts were made during the 1980s to amplify and correct the production files, identify and analyze all defects irrespective of their importance, and carry out regular tests on equipment sampled from the production line. The severity of these tests frequently exceeded the levels specified for the qualification tests. The flight data measurement plan that covered more than 700 parameters for which results were transmitted to the ground during the qualification flights—this figure was reduced to 400 for operational flights—made it possible to analyze each flight in the finest detail and thus acquire in-depth knowledge of the launcher. This unprecedented quality construction program, conducted by a European team highly motivated at all levels, meant that the 90% reliability objective assigned to the Ariane in 1973 was quickly exceeded. By the end of October 1999, a total of 118 Ariane launchers (versions 1 to 4 inclusive) had flown, and the reliability figure for Ariane 4 exceeded 97% after a string of 48 consecutive successful flights.
Marketing
Ariane marketing operations achieved rapid success, despite a number of technical difficulties encountered at the start of the program. The success of the development plan, combined with the appraisals conducted by potential customers, bred a high degree of confidence in the European launcher and the French Guiana ground facilities. The first non-European customer to place an order for Ariane launch services in 1978 was Intelsat (International Telecommunications Satellite organization), well known in the satellite world for its technical expertise. This choice was extremely important for Ariane and induced confidence on the part of other customers, who then decided to regard the Ariane launch system objectively. Apart from its own inherent merits, Ariane had the advantage of a favorable market situation. Commissioning of the American Space Shuttle led NASA to stop producing its conventional Delta and Atlas-Centaur launchers. Delays with the Shuttle, due to technical difficulties, discouraged satellite operators, obliging them to seek other launch possibilities. Furthermore, the high U.S. dollar exchange rate during the early 1980s made Ariane prices extremely attractive compared with the American launch systems. This marketing success grew further with the passage of time. The first Ariane 3 flight in August 1984 qualified the dual-launch capability of two Delta class spacecraft and significantly increased both the launch capacity and competitiveness of the Ariane system.
In 1985, Ariane launched the same number of commercial satellites as the Space Shuttle, and in the following year, Arianespace signed no fewer than 16 launch contracts. By 1987, Arianespace had 57 contracts (satellites to be launched or already launched) from the European Space Agency, the member countries of the European space community, international organizations, national organizations in non-European countries and private companies, representing a total sales value exceeding FF 16 billion (about $2.25 billion). Arianespace had acquired a share of more than 50% of the open satellite market, a position that the company has succeeded in maintaining despite increasingly severe competition.
Ariane Launch Site No. 2 (ELA 2)
France decided to construct a launch range in French Guiana in April 1964, after the deciding to discontinue launch operations from the Hammaguir range in southern Algeria. The French Guiana site was chosen as a result of a comparative study of a number of possible locations. Paradoxically, if we compare the 1964 situation with that of today, equatorial launches, if they were even mentioned, did not initially constitute a priority criterion. However, this view was quick to change, and in July 1966, the European launcher organization accepted the French proposal to contribute to constructing an equatorial launch base in French Guiana, initially intended for Europa 2. It should be noted that at that time, 33 years ahead of the Sea Launch project, the French Guiana site was in competition with a floating marine platform project proposed by Italy, similar to the San Marco platform which it was then operating offshore from Kenya.
A decision was quickly made in favor of the Kourou site (the population of the village of Kourou was 660 in 1964) on the Atlantic seaboard 70 km northwest of Cayenne in a sparsely populated region. The exceptional geographical characteristics of this site combined a wide launch arc over the ocean and favorable climatic conditions (infrequent storms, no cyclonic or seismic activity, and temperatures varying only slightly round a mean figure of 25°C). All types of mission (polar and equatorial orbits in the range — 10.5° to + 93.5°) could be planned in complete safety, an advantage that no other operational base possessed then. The closeness of Kourou to the equator (5°20 N) is ideal for placing telecommunications satellites into geostationary orbit. At this latitude, the slingshot effect induced by the rotation of Earth is near its maximum, and propellant consumption for adjusting the plane of the geostationary orbit is minimum. In global terms, the mass gain achieved with a launch from Kourou is approximately 17% compared with a launch from the Cape Canaveral, using an identical launcher.
The Guiana Space Centre (CSG) facilities were first operationally tested in April 1968 with the launch of a Veronique sounding rocket, and the Centre was inaugurated officially in 1969. The first satellite was launched on 10 March 1970 with the French Diamant B launcher. The first Europa 2 flight was in November 1971, using the new CSG launch pad. The failure of this launch had a series of consequences already mentioned at the beginning of this article. For both economic and political reasons, it was essential for the Ariane 1 launcher to make the best possible use of the heavy investments made in Europa 2. Nevertheless, certain facilities were unsuitable for launch operations with Ariane because of insufficient capacity and lack of operational flexibility and due to an absence of any growth potential. This made these facilities largely incompatible with the Ariane performance enhancements already under study. After proposals from the French Space Agency (CNES), the European Space Agency (ESA) decided to construct Ariane launch site No. 2 (ELA 2) in July 1980. This new facility was required to meet the following specifications:
• provision for 10 launches per year (Ariane 2, 3, or 4) and execution of two launches within an interval of less than one month;
• provision for replacing a launcher already on the pad, in case of need, by the launcher scheduled for the next flight;
• provision for preparing larger payloads with improved facilities for payload/ launcher integration.
The design of the new launch site differed from that of ELA 1. The time spent by the launcher on the pad had to be reduced to a minimum to allow for executing launch operations and post launch rehabilitation of the pad within a maximum 1 month. This assumed locating a launcher preparation zone on the edge of the safety perimeter.
Following their arrival from Europe, the stages are erected, assembled, and tested in the rear preparation zone. Then, the launcher on its mobile launch table is transferred to the pad via a dual rail track (Fig. 5). There it is connected to the propellant and fuel circuits and undergoes a launch rehearsal procedure, including filling of the third stage with liquid hydrogen and liquid oxygen to check the absence of ground and onboard leakage. The success of this operation authorizes assembly of the upper part (payloads, adaptors, and fairing were integrated beforehand in a dedicated building) with the launcher. The principles of the launch sequence qualified on the ELA 1 site have been retained. The ELA 2 pad uses many of the support facilities previously used for ELA 1.
Figure 5. ELA 1 and ELA 2 launch sites. ELA 1 site (foreground): Ariane 3 launcher during third-stage fueling tests on day D — 9. At rear: Ariane 4 launcher recently arrived on the ELA 2 pad. Note the double rail track between the pad and the preparation zone (background). The white circle midway down the rail track is a turntable, used to allow two launchers to pass each other if necessary.
Ariane 4 Program
Ariane 3 represented a short-term response that enabled the Ariane launcher to position rapidly vis a vis its two American competitors, Thor Delta and Atlas-Centaur. The arrival of the Space Shuttle at the end of the 70s and the policy adopted by NASA of marketing Shuttle flights at extremely attractive prices required a more comprehensive response from the Ariane program.
It was obvious that the NASA price policy would lead to a profound change in satellite design, broadly on the following lines:
• increase in Delta class satellite size and mass, up to the point of occupying, in a vertical position, the maximum volume available in the Shuttle cargo bay. This led to satellites of between 1400 and 1500 kg for injection into geostationary transfer orbit;
• appearance of a family of satellites installed in a longitudinal position in the Shuttle cargo bay, designed to take advantage of the large diameter available and reduce launch cost (dependent on the height of the satellite). Corresponding changes, that involved problems of both mass and diameter were more difficult to predict.
As regards satellite mass, there were some projects for TV satellites of 2500 kg. However, these were rare and still only moderately credible at the end of 1980, in particular among those who predicted a reduction in satellite mass values from miniaturization of satellite electronics. However, things moved in 1981, and the objective of 2500 kg was finally adopted.
The diameter problem was even more delicate. The Ariane program team foresaw major aerodynamic problems if the fairing diameter exceeded by too great an extent the 2.6 m diameter of the third stage on which it was mounted. However, determination of this value required aerodynamic tests that were too lengthy to conduct before making the decision. The diameter of 3.65 m finally available for payloads was in fact the result of a compromise between what the program team considered possible at the least risk and the sacrifice which satellite customers were prepared to accept for the advantage of a second launch service source.
Competitiveness requirements dictated the pursuit of the dual-launch policy implemented for Ariane 3. The GTO performance objective assigned to Ariane 4 was consequently an ability to execute simultaneous orbit injection of two payloads, one of 1400 kg and the other of 2500 kg, giving a total of 4300 kg, including the mass of the dual-launch structure. (Note that this objective had increased from 3400 to 4300 kg in 1981 and that the maximum performance demonstrated in 1999 exceeded 4900 kg.)
This program was to be regarded as a continuation of the Ariane family, in other words as a complementary performance enhancement phase, taking full advantage of work carried out for Ariane 2 and 3, and innovating as little as possible in propulsion. This was decided to take the fullest advantage of experience already acquired and to improve the reliability target from 0.90 to 0.92. Given the ambitious performance objective that had been set, the configuration adopted was required to embody a certain degree of flexibility with regard to lower mass payloads to apply an attractive pricing structure for a wide range of satellites. Single launches of heavier satellites like Intelsat 6—3600 kg—had to be possible, of course.
These performance objectives were achieved on the basis of Ariane 3 by introducing the following modifications:
— First stage propellant load increased from 147 to 227 tons, while retaining the same operating point for the Viking engines as qualified for the Ariane 3 program. Stage 1 burn-time was increased to 205 s.
— Development of a liquid propellant booster, corresponding to a reduced scale copy of the first stage as it uses the same engine with an appropriately adapted pressure ratio, and the same tank pressurization system. Identical stainless steel tanks carry 39 tons of propellant.
— Adaptation of the Ariane 3 solid propellant boosters. Carrying 9.5 tons of propellant, these boosters burn for 36 s and deliver 650 kN thrust each.
— Development of a new water tank, located on top of the UDMH tank. This tank is constructed in composite materials, and supplies the first stage and liquid propellant boosters in blow-down mode.
— Modification of the vehicle equipment bay structure, to achieve better pay-load integration flexibility and easier transition to the new fairing diameter.
— Development of an external Ariane dual-launch carrier structure (SPEL-DA), providing for dual launches of large-diameter satellites.
— Development of a new fairing with a useful diameter of 3.65 m.
This program was proposed by France to the European Space Agency and formally accepted in January 1982. The same program management principles were adopted as those for Ariane 1 and 3.
Launcher Development. The development phase did not introduce any major innovations in propulsion. First, the first-stage propulsion bay had performed extremely satisfactorily on the test bed in the Ariane 3 development program for burn-times very close to those required for Ariane 4. Furthermore, the booster propulsion system was based on a configuration already used to develop the Viking engine. Consequently, insofar as propulsion was concerned, the task was one of optimization or of demonstrating new operating margins because the burn-time for the Viking engine had been increased substantially. On the other hand, work relating to the launcher system was substantially more complicated than first thought. It was found necessary to adopt the complete system approach for each lower composite configuration. Furthermore, the increase in the height of the launcher and the size of the upper composite were reflected in a considerable increase in general loads that required substantial modification of a number of structures, including connecting flanges, in particular. Furthermore, first-stage in-flight stability was a major source of concern during the first few years of the program. Digital control was adopted to introduce a large degree of flexibility in the development time schedule and also with actual operation of the launcher by making it possible to finalize the configuration of each launcher only 2 months before lift-off. Today, Ariane 4 can place from 2.1 to 4.9 tons in a typical GTO orbit, depending on the number of liquid (L) or solid (P) propellant boosters fitted on the first stage (AR 40,42P,42L,44P,44LP,44L). Nine different configurations of SYLDA, SPELDA, and fairings were available for payload accommodation. The Ariane 4 maiden flight, initially planned for late 1985, took place successfully on 15 June 1988.
The Birth of Ariane 5
Initial reflection at CNES on Ariane 5 dates back to 1978. At that time, the Ariane 5 launcher was regarded more as a means to access low-orbit, manned-flight missions. Nevertheless, the launching of application satellites into geostationary transfer orbit continued as an objective, in particular in triple-launch mode, to pursue the Ariane scale effect competitiveness approach. The two missions rapidly acquired identical importance.
In contrast to the approach adopted for the American Space Shuttle, the primary principle adopted was that priority should be given to competitiveness for commercial flights. Manned spaceflight induces substantial costs that it is absurd to impose on operations that can be conducted by an unmanned vehicle. The system proposed had to be capable of manned flight and unmanned flight missions without increasing the cost of the latter. Consequently, the new launcher had to be regarded and optimized as a single entity at the design stage to ensure optimum use of available resources, and the manned flight aspect had to benefit from lessons learnt from unmanned missions, while also providing for substantially dissociated use when the operational phase was reached. The manned module was consequently designed as a ”special payload” mounted on top of the launcher. The capsule and spaceplane concepts were then analyzed in parallel. The latter presented the advantage of a considerable reentry cross-range and a high degree of orbit return flexibility that provided for a soft landing and consequent reuse. The onboard intelligence provided launcher ”brain” functions during the ascent phase and thus eliminated the need for a vehicle equipment bay required for unmanned flight. This, then, was the Ariane 5 – Hermes concept at the end of 1979.
At that time, Ariane 5 comprised the Ariane 4 first stage, and a new second stage burning liquid hydrogen and liquid oxygen (H 55), carrying Hermes for manned missions or the Ariane 4 third stage (H 10), the vehicle equipment bay, and fairing for unmanned flight. This study, which emphasized the need for a large cryogenic engine for future Ariane improvements, made it possible to propose and obtain funding for a 3-year French national program in 1980 and to commence work on the basis of design for an engine to deliver 600 kN thrust. Germany, Sweden, and Belgium joined this project under the terms of bilateral agreements. Partially Reusable Concept Studies. In 1980, the impact of the Space
Shuttle and its price policy led to strong criticism of the conservative characteristics of the Ariane 5 configuration. As a matter of principle, reusable concepts were regarded as more economical, the more so because the maintenance costs for such systems were ignored and operating costs were substantially underestimated. Considerable attention was then paid during that year to improving the cost levels and analyzing the operating costs announced for the Space Shuttle and those observed for the early stages of the Ariane operational phase. In parallel, the problems induced by the rehabilitation of equipment used during the Ariane 1 development phase demonstrated the importance of the corresponding work and the complementary qualification cost that would be required. A number of concepts were considered:
* recovery of the first stage by parachutes. This was tried unsuccessfully with the first stage of the Ariane launcher used for flight 14. The liquid propellant boosters for Ariane 4 were also initially designed to facilitate this type of recovery.
* design of a first stage comprising a stack of several H 55 stages, recovered individually using a delta wing;
* consideration of a winged first stage, etc.
This line of approach failed to produce any attractive, realistic solutions, and was abandoned early in 1982.
Comparative Configuration Studies. A systematic review of all possible concepts that matched the performance objectives was then undertaken. More than 24 different configurations were drawn and assessed for performance and cost. A configuration involving a large cryogenic core stage flanked by two large solid propellant boosters was examined and then abandoned because it was impossible to identify a derivative solution, with reduced performance, which would have made it possible to adapt a launcher configuration to the mission model in the same way as with Ariane 4. This configuration was looked at again, when the desired degree of flexibility had been introduced by offering the choice of an Ariane 4 third stage (H 10), or alternatively a highly simplified storable propel-lant stage, and after further studies had made it possible to set an acceptable price objective for manufacturing the solid propellant.
Finally, these three configurations were selected for more detailed comparative analysis in mid-1983:
* a solution that corresponded to direct continuity with Ariane 4 but introduced a cryogenic second stage. This solution was derived directly from experience acquired from Ariane 4 and made it possible to evaluate cost and performance extremely precisely.
* a solution based on the above, replacing the Ariane 4 first stage with a cryogenic stage equipped with four or five engines identical to that used for the second stage. This solution had the advantage of requiring only development of a single propulsion engine, namely a high-thrust LOX/LH2 engine.
* a solution involving a large cryogenic core stage, flanked by two large solid propellant boosters. This basic composite carried a simple storable propellant upper stage for low-performance missions or a cryogenic stage that carried 10 tons of LOX/LH2 (Ariane 4 third stage) for high-performance missions.
All three configurations required developing of a new LOX/LH2 engine This conclusion led France to propose the development of an engine of this type, based on the project commenced in 1980, to the European Space Agency. A further year was then devoted to a detailed comparative study of the three configurations.
The configuration that had large solid propellant boosters was finally selected on the basis of intrinsic reliability, recurrent cost, greater potential, and development time schedule control criteria. Only the safety criterion remained questionable. Compared with the abrupt failure of solid propellant boosters, their liquid propellant counterparts are always presented as easy to control. Nevertheless, in-depth analysis of the complete system demonstrates that this is not always the case, and that even in this favorable situation, there are still flight phases where the spaceplane (in contrast to a capsule) simply cannot accomplish its true rescue mission. (Long after this choice had been made, this frequently impassioned discussion was stimulated once more following the Challenger failure.)
Therefore, the configuration that incorporated two solid propellant boosters was finally presented in 1985 and adopted as a European Space Agency program in 1986 and 1987. Further optimization led to incorporating the following aspects:
• abandonment of the triple-launch concept in a basic mission context due to the excessive operational constraints relating to the satellites (availability);
• increase in payload mass values: the nominal performance objective was then dual launches into geostationary transfer orbit for satellites of 2950 kg each;
• mass problems encountered with the Hermes program. This led to an increase in the performance of the lower composite and abandonment of the cryogenic upper stage.
Ariane 5 Specifications. Final specifications for the Ariane 5 program were as follows:
* simultaneous launch of two satellites of 2950 kg and a diameter of 4.53 m into geostationary transfer orbit under environmental conditions and with a degree of precision, etc., comparable with Ariane 4 launches, representing a single-launch performance equivalent of 6800 kg;
* launch of an 18-ton payload into a circular orbit of 550 km, inclined at 28°5′;
* Hermes launch: this mission was not to introduce constraints liable to penalize unmanned flight. Consequently, performance requirements for this mission were deduced from those specified for unmanned flight, taking due account of safety constraints;
* Reliability of unmanned flight was set at 0.98, almost 10 times higher than the initial specification for Ariane 1. This ambitious target was justified by the high cost of insurance for both launcher and satellites (over 20% in 1988) and the major consequences of a flight failure. The safety of the crew was consequently provided forby ejection of the Hermes spaceplane on detection of an operating anomaly. The safety objective for the crew was set at a figure of 1 – 10 – 3.
* Cost objective per launch, on the basis of eight launches per year, was set at 90% of the figure for an Ariane 44 L. Given the respective performance of the two launchers, this corresponded to a reduction of 45% in the cost per kilogram in orbit for an equivalent fill factor.
• The target for a maiden flight was initially set for April 1995. It was subsequently put back by one year due to a number of economic constraints.
As for earlier Ariane programs, management was delegated to the French Space Agency (CNES) by the European Space Agency. The conditions for this delegation of authority were more restrictive in this case, however, due to the need for close coordination with the Hermes program and the now powerful image of the Ariane program.
The three main management principles stated at the beginning of this article with regard to Ariane 1 were retained and in fact strengthened in two areas:
* Management specifications incorporated a design-to-cost objective.
• The safety/reliability approach was further emphasized through systematic integration of past experience, both from incidents and accidents that occurred during testing, in the course of previous Ariane flights, or in the context of other programs (Challenger, etc.).
Furthermore, faced with an identified failure mode, the principle of a dual approach, involving simultaneous reduction of the probability of that failure and improvement of system tolerance was imposed.
Ariane 5 Launcher
The Ariane 5 launcher stands approximately 51. 5 m high; the actual figure depends on the upper composite configuration, and the lift-off mass is 740 tons. Lift-off thrust is 11,660 kN. On the ground, the central core vehicle that has an outside diameter of 5.46 m is suspended at the level of the first-stage forward skirt between two solid propellant boosters. This connection, through which the thrust of each booster is introduced in the core stage, is made of alternate elastomer and metallic shims, designed to provide a damping effect on vibrations induced by booster combustion (Fig. 6).
The cryogenic main stage, developed under Aerospatiale as the prime contractor, is 30.5 m, and its dry mass is 12.2 tons. This stage contains 158 tons of liquid hydrogen and liquid oxygen. The Vulcain engine (Fig. 7) is mounted on a thrust cone that distributes thrust evenly at the base of the liquid hydrogen tank. The engine can be oriented along two orthogonal axes by hydraulic actuators operating on the lost fluid principle. This fluid is stored in tanks operating in the blow down mode. Connecting struts between the main stage and the solid propellant boosters provide rigidity for the rear part. The light alloy propellant tanks have a common bulkhead, insulated with expanded polyurethane. The hydrogen tank has a volume of 390 m3 and is pressurized at values that vary according to the flight phase (between 2.15 and 2.35 bar), using hydrogen tapped at the outlet from the engine regenerative circuit. The oxygen tank (120 m3)isin the upper position and is pressurized with helium gas (3.5 dropping to 2.85 bar), obtained by heating liquid helium in a heat exchanger located in the oxygen turbine exhaust line. This helium (1.15 m3) is stored in a super insulated tank mounted on the thrust frame. The engine control system is supplied with helium gas stored under pressure in separate tanks. A thrust frame, secured to the upper part of the oxygen tank, receives the thrust of the solid propellant boosters, which is then distributed evenly toward the upper composite.
Figure 6. This is a cutaway drawing of Ariane 5. The core is the Vulcain Engine with a payload of two satellites shown above the fuel tanks. The two solid strap-on motors are shown on either side of the core.
Figure 7. ARIANE 5—Vulcain flow diagram. The Vulcain engine delivers 1140 kN thrust in vacuum and has a specific impulse of 432 sec. It uses conventional gas generator cycle technology. The propellants are delivered by two independent turbopumps. The liquid hydrogen unit operates at 34,000 rpm and comprises a two-stage centrifugal pump, preceded by an inducer that ensures favorable intake characteristics. This pump is driven by a 12-MW, two-stage turbine. The pump delivers up to 560 L/s of liquid hydrogen at a pressure of 17 MPa. The single-stage liquid oxygen turbopump operates at 13,400 rpm and delivers 177 L/s of propellant at 13 MPa (3.7 MW). The two turbines are driven in parallel by a single radial injection gas generator, that operates at a combustion pressure of 8 MPa. The generator is supplied with propellant tapped off at the pump outlet. The combustion chamber pressure is 110 bar. The liquid hydrogen enters the propulsion chamber via an annular distributor. Most of this flow is routed through channels integrated in the double-walled structure of the combustion chamber and throat assembly. The nozzle is cooled by a simple process known as dump cooling: the remaining hydrogen flow is routed through 460 spirally welded iconel tubes, whose diameters increase to give a continuous, bell-shaped surface, then escapes through micronozzles set along the bottom rim of the main nozzle. Although these gases do not undergo combustion, they are heated during the trip and contribute to overall thrust. The turbopumps, gas generator, and combustion chamber ignitions are started by pyrotechnic cartridges. The mixture ratio (mean value 5.25) is adjusted by a two-way valve, used to modify this ratio to terminate combustion on quasi-simultaneous depletion of the two propellants.
Europropulsion (joint subsidiary of SNECMA/SEP and FIAT/BPD) was the prime contractor for development of the Ariane 5 solid propellant boosters.
Each booster is 31.16 m high, with a diameter of 3.05 m and a post-combustion mass of 39.3 tons. It contains 238 tons of solid propellant grain, a composite with an ammonium perchlorate and polybutadiene base charged with aluminum. The booster casing is made of a high-strength low-alloy carbon steel and comprises seven cylindrical sections and two bulkheads. The sections are flowturned to a thickness of 8 mm from forged preforms and then assembled using a tang and clevis connection. The sections and bulkheads are assembled to form three segments, each of which is loaded independently with propellant. Internal thermal insulation, made of rubber-based, silica or fiber-filled material, protects the structure from hot combustion gasses. The forward segment is loaded with a 20 tons, star-shaped solid propellant block in Italy. In view of the mass and size of the boosters, a dedicated plant has been constructed in French Guyana for fuelling the central and rear segments (approximately 110 tons of pro-pellant each).
The nozzle, with a flexible bearing made of alternate elastomer and metallic shims, can be steered up to 6° to control the thrust vector. The hydraulic actuators are driven by fluid stored within high pressurized, carbon-fiber vessels operating in blow-down. This fluid is ejected at the nozzle exit. The nozzle, which is highly integrated with the motor, represents a prudent extrapolation of the nozzles developed and qualified for defense applications. The throat in carbon/ carbon material ensures minimum erosion during flight. The exit cone is composed of a light alloy housing, with phenolic carbon and silica insulation.
The boosters are jettisoned from the core stage by means of pyrotechnic cords. Solid fuel thrusters located in the rear part and in the forward cone are used to distance the boosters from the launcher in a radial plane.
The booster thrust evolves in flight in order to limit general loads, with a maximum of 6,700 kN and an average value of 4,900 kN in vacuum. It must remain within very tight tolerances to limit any thrust differential between the two boosters at any moment. The maximum combustion pressure is 60 bar, and the specific impulse in vacuum is 270 s.
The vehicle equipment bay is 1.56 m high and constitutes a linkage structure between the first stage, second stage, and fairing.
Developed under the prime contractorship of DASA, the second stage is fitted inside the vehicle equipment bay. This is an internal and relatively compact stage, with a diameter of only 3.94 m and a height of 3.36 m. The dry mass is about 1,250 kg. This second stage carries the payload adaptor (Fig. 8).
The propulsion system comprises an “Aestus” engine, burning storable hypergolic propellants (MMH and N204), loaded in helium-pressurized tanks and consequently requiring no turbopump. This technological solution was adopted for its reliability, and its simple operation and re-ignition. A total of 6,550 kg of N204 and 3,200 kg of MMH (maximum load mass) are contained in two pairs of identical cylindrico-spherical tanks, arranged axisymmetrically two by two. The lift-off mass is 11 tons.
The engine combustion chamber, extended by a nozzle in refractory steel, is cooled by an MMH circuit. The mixture ratio is adjusted to a value of 2.05 by calibration on acceptance testing. The engine can be oriented through an arc of +4.8° on two axes, using electrical servo actuators. The Aestus engine develops 29.1 kN thrust and has a specific impulse of 324 s. The burn-time is about 19 min.
Figure 8. Ariane 5 second stage.
Other details of Aestus engine are as follows:
Propellant feed rate: 8.77
Nozzle expansion ratio: 84
Mass: 111 kg.
The bay carries all functional electrical equipment used throughout the mission and a hydrazine attitude control system operating in the “blowdown” mode. This system is used to control roll throughout the propulsive flight, except during the solid booster burn phase, and during the attitude control phase after shutdown of the second-stage engine. By comparison with the Ariane 1 to 4 programs, the functional electrical systems have been fully duplicated and operate in the full active redundancy mode. This might allow for progressive incorporation of standard electrical components.
The vehicle equipment bay is separated from the main stage by dilation of a sealed pyrotechnic tube, which causes a fragile flange to rupture. The upper composite is then distanced by pyrotechnic actuators.
The assembly sequence and facilities in French Guiana (Launch site N° 3: ELA3) are based totally on experience with Ariane 4, taking due account of the new mass and dimensional values involved, and carrying the policy of stripping the pad down to the bare minimum a step further.
As for the previous Ariane launchers, the synchronized sequence commences at time T — 6 min, 30 s. It is subdivided into two separate automatic procedures that involve the fluid systems, on the one hand, and the electrical systems, on the other. At T — 4.5 s, an execution command signal is sent to the onboard computer, which then activates the inertial guidance systems and authorizes ignition of the Vulcain engine at time T. Following verification of correct engine combustion at T + 6.2 s, the solid propellant boosters are ignited at T + 7 s. The boosters burn for about 130 s. They are jettisoned after burnout, when an acceleration limit of about 5.2 ms2 is detected. For a typical GTO launch, booster separation occurs at an altitude of approximately 60 km at about T +140 s. By this time the launcher relative speed has reached approximately 2000 m/s. The fairing is jettisoned on the basis of a thermal flux calculation, at about T + 3 min 10 s. Shutdown of the Vulcain engine is initiated by measuring propellant depletion at about T + 9 min 46 s. Launcher speed is then approximately 7600 m/s. After separation, fallback of the main stage is controlled so that the stage hits the ocean approximately 1800 km from the Colombian coast. The upper stage is ignited at about T + 10 minutes. During the launch, onboard telemetry is transmitted to Kourou via ground stations located at Natal (Brazil), Ascension Island, Libreville (Gabon), Hartebeesthoek (South Africa) and Mali-ndi (Kenya). The upper stage is fully passivated after separation of the payloads to minimize orbital pollution.
The first Ariane 5 flight was a failure due to a software fault. The second flight (502) took place on 3 October 1997. During the second flight, the only anomaly observed was an excessive roll torque at the limit of acceptability, following jettisoning of the solid propellant boosters and during the Vulcain engine burn phase. This fault was corrected very simply by a minor modification to the orientation of the turbine exhaust nozzles.
The third flight demonstrated the exceptional flexibility of the launcher, based on the simultaneous execution of two largely different missions. The first involved placing a reentry capsule into a suborbital trajectory. The precision of the point of impact in the Pacific was excellent. The other mission involved injecting a payload into a geostationary transfer orbit, again achieved with excellent precision and followed by a second-stage maneuver. This mission involved several successive reignitions and shutdowns of the second stage.
Comparison Between Ariane 44 L And Ariane 5
Unquestionably, Ariane 5 a technical break with earlier members of the Ariane family. However, continuity in human resources, management principles, and industrial organization ensured that the Ariane 5 program benefited from all experience previously acquired and made it possible to improve the intrinsic reliability by one order of magnitude. The Ariane 5 launcher has four active stages, whereas Ariane 44 L has seven. The number of separation sequences has thus been reduced. Ariane 5 does not have a launch table that is active in positive time, as is the case with Ariane 4. Nor does Ariane 5 have cryogenic fuelling arms that retract at the last moment before ignition and liftoff. Ariane 5 fluid connections are passive and are pulled free on lift-off. The aborted launch case, which requires extremely complex revalidation operations for Ariane 4, is thus simplified to the extreme. Vulcain engine ignition on the ground was deliberately chosen (following ignition failures observed on Ariane 3 flights 15 and 18), as was also the decision to wait to establish a stationary regime for the complete stage (following the flight 36 failure) before checking the engine and authorizing ignition of the solid propellant boosters. The price of this option is a loss of performance (Vulcain engine thrust was set to ensure minimum acceleration of the launcher after boosters were jettisoned. On the other hand, the Vulcain nozzle could have been adapted more efficiently in the case of in-flight ignition). The production of hydraulic power required to operate the booster and first-stage attitude control actuators is obtained from simple pressurized tanks on Ariane 5, compared with the hot-gas-fed engines used for Ar-iane 4. This list, which covers only simplifications visible at launcher system level, can be extended for each subsystem and component.
Ariane 5 and Hermes
It has frequently been said or written that the Hermes project led to the design of a launcher which was too large, or nonoptimum, in stage configuration. The need to improve performance yet again, to face up to the competition and remain competitive, clearly demonstrates that the first criticism was foundless.
No choice was made to the detriment of optimization for GTO launches. On the other hand, where a number of possibilities were equivalent for this optimization, choices were based on a Hermes criterion. Speaking in 1999, one can be thankful that this approach was adopted for at least three reasons:
• The Ariane 5 launcher is fully adapted for all types of missions, from constellations to heavy GEO spacecraft and servicing of the Space Station. The 503 flight demonstrated a high degree of flexibility during a single mission.
• The growth potential of Ariane 5 is very substantial, including replacement of the existing upper stage by a cryogenic stage.
• Manned spaceflight represented an excellent stimulus for the construction of launcher reliability. In this area again, the approach adopted was in no way unfavorable for unmanned missions. Abandonment of the Hermes program changed none of the choices made before this decision.
Ariane 5 Evolution Program
The increase in individual mass values of satellites in the Arianespace order topic continued steadily throughout the 1980s. Extrapolation of this growth demonstrated that Ariane 5 performance was likely to be insufficient by about one ton shortly after qualification of the new launcher. However, this demonstration failed to convince at the beginning of the 1990s. The Hermes spaceplane program was encountering difficulties, including mass-related problems, and the proposal to increase the performance of the Ariane 5 lower composite was consequently regarded as a subterfuge for increasing the size of a launcher, still regarded as too big for launching commercial satellites. Furthermore, the multiplication of constellation projects involving small satellites in LEO and the increased credibility of plasma propulsion techniques for satellite orbit control provided ammunition for those who supported stopping further increases in the mass of geostationary satellites. When the decision to go ahead with the proposed enhancement program was finally made in 1995, this was based more on the desire to preserve cryogenic propulsion expertise than on a genuine need to enhance Ariane 5 performance, even with no change in production costs. The new enhancement program, designated “Ariane 5 Evolution” is aimed at achieving capacity for simultaneous launch of two 3300-kg-class satellites into geostationary transfer orbit.
The following modifications were adopted to meet this specification:
* Increase in Vulcain engine thrust from 1140 to 1350 kN, while changing the mixture ratio from 5.3 to 6, but without modifying the external dimensions of the first-stage tank (although the position of the intermediate bulkhead changed). First-stage mass is increased up to 170 tons of liquid hydrogen and liquid oxygen.
* To recover the specific impulse lost through this modification of the mixture ratio, the turbine exhaust gases are injected into the nozzle, whose expansion ratio area is increased. This technical solution, not adopted for the basic Vulcain program to avoid excessive integration of elements supplied by different contractors, was validated by a technological program in parallel with the main development program.
• Replacement of connections within the three main segments of the solid propellant boosters by welds. In addition to a substantial booster mass gain and lower costs, this simplifies application of internal thermal protection.
* Adaptation of the dual-launch system (SYLDA) used with Ariane 2, 3, and 4 to the dimensions of Ariane 5.
In parallel with the Ariane 5 Evolution program, Arianespace is proceeding with a number of optimization actions (lightening of the vehicle equipment bay, increasing the solid propellant loading of the booster forward segment) based on experience acquired during the development phase. Combined with Ariane 5 Evolution, this will increase the performance in GTO up to 8 tons in single-launch configuration. These enhancements should be operational late mid 2002.
New Phase: Ariane 5 Plus
Changes in the launch service supply situation due to the upcoming availability of Atlas 3, Delta 3, and the American EELVs (Delta 4 and Atlas 5) will further intensify competition and induce a new trend in satellite mass values toward 5 tons for multimedia spacecraft.
Figure 9. The Ariane rocket family.
To preserve the competitive advantage achieved with dual launches, the performance of Ariane 5 must exceed 10 tons for GTO injection. This objective will be achieved in two phases:
• For mid 2002 horizon, a second stage will be developed with a propellant load of 14 tons LH2 and LOX, using the current Ariane 4 third-stage propulsion system. This, combined with Ariane 5 evolution, will take payload capacity up to 10.4 tons in single-launch configuration.
* For a late 2005 horizon, the HM 7 engine will be replaced by the new Vinci motor that delivers 15 tons of thrust and designed for repeated reignition. Tank capacity will be increased to 23 tons of LH2 and LOX.
This will provide geostationary transfer orbit performance of 11.9 tons in single-launch configuration.
The existing storable liquid propellant upper stage will also be retained but will be adapted to be capable of multiple ignition, with long ballistic phases, to have the flexibility required for all type of missions.
Phase 1 and 2 of this evolution program were decided in the European Space Agency Council in 1998 and 1999. Work on the new engine has been authorized so that the relevant decision can be made in 2001.
Conclusion
Figure 9 briefly summarizes the successive changes in the Ariane launcher and the policy of addressing each new phase takes fullest advantage of experience acquired in the preceding phases. Indeed, Ariane 5 shows a break in this process, and has more the air of the start of a new family. However, the actual break is substantially less pronounced than appears to be the case because the same teams, strengthened by working together for 20 years, have developed and will manufacture Ariane 5.
The technical solutions adopted have frequently been criticized for a lack of initiative because there is sometimes a tendency to confuse efficiency with the technologies applied. But a launcher provides a service, and the only thing that counts is the quality of service provided to the satellite owner. This service is judged in performance, operating cost, reliability, availability, appropriateness of satellite preparation facilities, etc. These are the criteria on which the program’s efforts primarily were focused.
