Description of the FESTIP VTHL-TSTO System Concept Studies
Martin Bayer
1. General Background of FESTIP
FESTIP was an ESA program to establish an approach to reusable launchers and to prepare the associated technological basis. Specific objectives were the investigation and evaluation of promising concepts for potential future reusable space transportation systems, the derivation of the respective technology needs and the definition of associated technology development and verification plans. The program started in 1994 and continued until the end of 1998. It comprised a system study as the focal point of the activities and technology studies in the areas of structures, materials, rocket and airbreathing propulsion, aerothermodynamics and heat management.
The system study was led by DASA, with about 30 European companies from Austria, Belgium, Germany, Italy, the Netherlands, Norway, Spain and Sweden participating as subcontractors. It was performed by an international integrated system concepts team and based on commonly agreed mission scenarios, system requirements, design standards, analysis tools, technology assumptions and assessment criteria. A variety of SSTO and TSTO concepts with different launch and landing modes and pure rocket as well as airbreathing propulsion were considered. The definition of the concept alternatives was based on previous national European activities and projects and various recent RLV proposals from the USA as well as a series of related ESA studies, such as the WLC studies.
Some of the concepts analyzed in FESTIP are shown in Figure 1, such as the rocket propelled VTHL-TSTO FSSC-9, the HTHL-TSTO FSSC-12 with an airbreathing lower stage, which was derived from the German SÄNGER concept, and the so-called Hopper suborbital rocket propelled HTHL-SSTO FSSC-15 with an expendable kick stage, which is based on the German ASTROS design. All vehicles were designed for cryogenic propulsion.
This report describes the design heritage and evolution of the rocket powered VTHL- TSTO, as well as its main characteristics and the respective results achieved in FESTIP.
1 Figure 1: Various Winged RLV Concepts Studied in FESTIP
2 2. FESTIP System Requirements and Design Standards
The main technical, economic and programmatic assumptions, ground rules and system requirements, which had to be met by the various FESTIP concepts, are as follows:
• Layout as a reusable launch vehicle
• Reduction of the specific transportation cost to 30% of the present respective value of an expendable launch vehicle, such as Ariane 5
• Technology readiness target date of 2005
• Initial operational capability target date of 2015
• Maintained system lifetime of 30 years
• Fleet size of 3 vehicles (2 operational, 1 in maintenance cycle)
• Nominal yearly launch rate of 24 missions
• Maximum yearly launch rate of 36 missions
• Unmanned transport of 15,432 lb into a circular orbit of 135 NM/5° and return with payload to Earth in an abort situation (Design Mission 1)
• Unmanned transport of 4,409 lb into a circular orbit of 135 NM/98° and return with payload to Earth in an abort situation (Design Mission 2)
• Usable payload bay envelope with 15 ft diameter and 32.8 ft length
• Launch from and nominal return to the Centre Spatial Guyanais in Kourou, French Guyana
• Evolution capability to unmanned payload retrieval and RVD (2nd generation)
• Evolution capability to manned missions and operations (3rd generation)
• Mass margin of 14% on structural components of orbital stages and of 12% on structural components of booster stages
• Mass margin of 12% on rocket propulsion
• Mass margin of 10% on subsystems
• Single engine out abort capability over the entire powered trajectory
• OMS ∆v capability of 459 ft/s
3 3. Origins and Evolution of the FESTIP VTHL-TSTO Concept
The VTHL-TSTO FSSC-91 with parallel staging and cryogenic rocket engines, which constituted the point of departure for the FESTIP activities on this vehicle type, was an adaptation of the German EARL II2 concept to FESTIP system requirements, design standards and technology levels. This design had in turn been synthesized from the earlier German ADV and EARL3 RLV studies. The EARL and EARL II designs are shown in Figures 2 and 3. Both semi- and fully reusable versions were considered for the different variants of the rocket propelled VTHL-TSTO, with the semi-reusables combining the winged boosters with an Ariane 5 core stage. Figure 4 shows the fully reusable FSSC-9 version.
Figure 2: EARL Concept Figure 3: EARL II Concept
Figure 4: Fully Reusable Version of FSSC-9
4 A detailed design activity on this concept was performed in the first phase of FESTIP. Booster and Orbiter are fully reusable and return to the launch site after mission completion. Table 1 contains the main technical data of the system, and Figure 5 shows an inboard profile.
Table 1: Main Data of FSSC-9
Stage Booster Orbiter Total Vehicle Length [ft] 153.5 128.9 Maximum Fuselage Width [ft] 22.3 18.4 Wing Span (including Winglets) [ft] 87.9 62.0 Usable Ascent Propellant Mass [lb] 727,072 178,351 Stage Dry Mass [lb] 115,741 66,137 Individual Stage Liftoff Mass [lb] 867,945 278,439
Figure 5: Inboard Profile of FSSC-9
The achievable payloads are 19,841 lb into the 135 NM/5° orbit for a launch mass of 1,146,384 lb respectively 5,732 lb into the 135 NM/98° orbit for a launch mass of 1,132,348 lb, which exceeds the requirements for both Design Missions. The liftoff acceleration is 1.4 g. Staging occurs at M 9.4 in an altitude of 36 NM. The return of the booster requires a turbojet propelled range of about 432 NM. The cross range of the orbiter is sufficient to glide back to the launch site from both target orbits.
The vehicles were based on simple geometries using mainly 2D-curved surfaces to facilitate scaling in the beginning of the study as well as to enable a low production cost
5 target. A fairly high nose radius was initially chosen for both stages to decrease thermal loads in that region, which was later reduced. It was however realized from the start that especially with respect to aerodynamic performance these shapes would have to be further refined for the final layout.
Both stages incorporate similar technologies for structures, propulsion and TPS. The primary cold structures and the LH2 tanks are CFRP, while the LO2 tanks are made out of Al-Li. Staged combustion cycle rocket engines with a chamber pressure of 3540 psia and a mixture ratio of 6.6 are used for the main propulsion. The low temperature TPS on both stages consists of a metallic multiwall concept, while for the higher temperature regions on the orbiter FEI, C/SiC shingles with IMI and C/C hot structures for the nosecap and leading edges are used.
The booster comprises a single integral LO2 tank in the nose, an integral LH2 tank in the middle and a small nonintegral LH2 tank in the rear. Ascent propulsion consists of four high performance main engines with fixed nozzles. Flyback of the booster is performed using two kerosene fueled turbojet cruise engines stored in the rear fuselage below the nonintegral LH2 tank and between the main landing gear. The engines are swung out downwards for operation.
The orbiter features an integral LH2 tank in the nose and a nonintegral LO2 tank in the rear. The payload bay is located between the two tanks close to the CoG. One single main engine of the same type as those on the booster, but with an extendible nozzle, is operated in parallel with the booster engines during ascent. Until staging, propellants are supplied from the booster to the orbiter via cross feeding. Two cryogenic OMS rocket engines are arranged on both sides of the main engine.
The configuration of FSSC-9 as a TSTO with parallel staging, which was based on the general EARL II layout as opposed to the tandem arrangement of the original EARL design, was critically reassessed in the course of the design process. Parallel staging and main engine burn with propellant crossfeed from the booster to the orbiter has repeatedly been traded off against other alternatives, such as parallel staging without crossfeed, tandem staging and staggered staging in a NASA study4 and also in the context of a French VTHL-TSTO design effort called TARANIS5, and has in both cases been selected as the most promising alternative with respect to performance as well as technical and operational aspects. The TARANIS concept, which is generally similar to
EARL II as well as the FSSC-9 except for the use of expendable LH2 tanks on the
6 orbiter, the lack of flyback propulsion on the booster due to a downrange transatlantic landing site and the belly on back arrangement of both stages, is shown in Figure 6.
Figure 6: TARANIS Concept
The tank arrangement and cross feed logic of FSSC-9 of connecting the tanks of the booster to those of the orbiter rather than directly to the orbiter engines and using gravity respectively the vehicle acceleration to support the transfer of the LO2 as the
heavier and denser propellant component is analogous to that for the cryogenic propellant tanks in both stages of the NASA concept. The NASA study identified no major potential problems for the viability of this solution, even in the context of three propellant components.
Several other deficiencies became however evident after reworking the EARL II design to conform to FESTIP requirements and design standards. During the FSSC-9 conceptual engineering and analysis, the following major problems and drawbacks were encountered:
• Technologically very demanding rocket engines had to be chosen due to geometric restrictions concerning the integration in the propulsion bay of the booster
• The deployment and retraction of a two position orbiter main engine nozzle introduced critical failure cases for both ascent and reentry
• In case of a failure of the single orbiter main engine no ascent abort capability was achievable due to lack of propulsion redundancy
• The orbiter propellant tank arrangement was not optimal with respect to ascent CoG migration considerations
7 • The ascent CoG shift was significant and led to the gimbaling requirements of the booster engines reaching the limit
• Potential aerodynamic wing interference effects of the two stages with flat undersides being mated belly to belly were unclear
• The rear solid separation motor locations on the booster led to plume impingement on the orbiter TPS
• High aerodynamic booster drag led to a significant propellant demand for the airbreathing engines used for flyback, which was initially underestimated
• The use of only two airbreathing engines for the powered booster flyback precluded an engine out capability during the return to launch site cruise flight
• The location and accommodation of the airbreathing engines necessitated the adoption of a very sophisticated landing gear on the booster
• The booster size chosen for the fully reusable system was overdimensioned for an application in a semi-reusable version with the Ariane 5 core stage
Despite these drawbacks the basic concept was still regarded as attractive, since a TSTO is consistently less sensitive against possible design deteriorations, like increases in the dry mass due to difficulties in the technology development, and thus exhibits the lower development risk than an SSTO. A TSTO also becomes feasible at a lower technology level.
Therefore, a complete revision of the concept was performed in the final phase of FESTIP, and the lessons learned from the analysis of this concept were integrated into the subsequent FSSC-16 design to obtain a more attractive TSTO configuration.
The goal was to create a concept family based on a reusable booster, which could be combined either with the Ariane 5 core in a semi-reusable version or with an orbiter for a fully reusable system. In addition to an advanced technology booster, a variant with state of the art propulsion based on the Vulcain engine used in the Ariane 5 core stage was explored as a possible intermediate development step for the semi-reusable application.
In order to decrease the technology risk associated with the rocket engines, it was decided to design the fully reusable version of FSSC-16 with a lower performance
8 rocket propulsion system based on staged combustion main engines with 2176 psia chamber pressure and fixed nozzles. The design features of parallel staging with crossfeed and a belly to belly arrangement were maintained from the FSSC-9 configuration. In the semi-reusable version, no crossfeed is performed.
The booster size was driven by the need of compatibility with the Ariane 5 core and the desire to maintain or even increase Ariane 5 payload capability in the advanced semi- reusable configuration, which determines the staging point and the required booster performance. A booster with a high staging Mach number corresponding to the current staging conditions requires either airbreathing or rocket propulsion for the return to the launch site, but the use of airbreathing engines enhances the operational flexibility and enables higher adaptability and evolution potential. To reduce overall system complexity, the fuel for the airbreathing and rocket engines of the booster should be the same. The use of Hydrogen as fuel for the airbreathers also reduces the propellant mass required for the flyback and consequently the overall launch mass and simplifies the launch operations. As a result, a booster layout with airbreathing engines using GH2 to enable the return to the launch site from a high staging Mach number was chosen.
Reusable winged cryogenic boosters for the Ariane 5, such as the concept shown in Figure 7, were studied before in the RRL6 study series of ESA.
In these studies, it was found that a configuration with two boosters, equipped with two Vulcain 2 engines derived from the original Ariane 5 core stage propulsion each, was economically unattractive, due to the low performance resulting from the thrust level achievable with four booster engines and the operational implications of having two boosters. Therefore, a design approach based on a single reusable booster with a higher thrust level was considered promising.
9 Figure 7: Ariane 5 with Twin Reusable Cryogenic Boosters
Consequently, a single winged booster with five Vulcain 2 engines and a propellant loading of about 430,000 lb, which was upscaled linearly with the number of engines from the value for the separate boosters, was taken as the point of departure for the initial sizing of the semi reusable version of FSSC-16. These data also took into account that the FSSC-9 booster had been found to be considerably oversized for an application as an Ariane 5 booster. For the advanced booster version used in the fully reusable configuration, the staged combustion engines with higher specific impulse, but the same thrust level as the Vulcain 2, would replace them.
Once the booster size had been established, the orbiter was sized accordingly to fulfil the FESTIP mission requirements for the fully reusable version. In this process, it was found that that a harmonization of the booster sizing with the Ariane 5 core as well as the orbiter allowed to arrive at the same shape and external dimensions for both winged stages, so that a common basic design appeared feasible. The development and production cost reduction potential associated with identical or similar stages had been pointed out in a comparative study of various RLV designs7. Therefore an ‘almost’
10 Siamese configuration with a high degree of commonality between booster and orbiter concerning engines, structures and subsystems was chosen.
Since the altitude/velocity flight envelope of the booster is completely contained within that of the orbiter, the design of both stages was derived and scaled from that of the orbiter of the HTHL-TSTO FSSC-12 shown in Figure 1, which was in turn based on the HORUS orbiter originally developed for the SÄNGER concept8 shown in Figure 8 and later refined in the WLC program9. This wing-body configuration incorporates near circular fuselage cross sections, conformal integral tanks in the front fuselage, double delta wings with winglets for directional stability and a single body flap.
Figure 8: SÄNGER Concept
The approach of adapting an existing layout was taken in order to make maximum use of available results and facilitate the design process. Due to the concurrent design activities on several concepts within FESTIP and the limited timeframe available, only one design loop could however be performed on this concept. As a result, several design details had to remain nonoptimized, but based on the insights gained in the design of FSSC-9 the following major changes were incorporated in the configuration:
• Use of completely identical lower performance rocket engines without extendible nozzles on both reusable stages
• Implementation of two main engines for redundancy on the orbiter
• Adoption of the FSSC-12 orbiter moldline for both stages
11 • Adoption of the FSSC-12 orbiter inboard profile for the orbiter
• Implementation of four Hydrogen fuelled turbofans for the booster flyback
4. Design Characteristics of the ‘Almost Siamese’ VTHL-TSTO Concept
The FSSC-16 represents a rocket powered VTHL-TSTO with a semi-reusable version, combining a reusable winged booster with an Ariane 5 core stage, and a fully reusable version consisting of the booster in connection with a winged orbiter. Both versions are shown in Figure 9.
Figure 9: FSSC-16 Fully and Semi-Reusable Versions
The general design of the fully reusable system was based on the principles of commonality and modularity. Both winged stages have the following common features and elements:
• Aerodynamic shape and geometric size
• Front integral LO2 tank and insulation
• Front integral LH2 tank and insulation
• Wing and winglet structure
12 • Aerodynamic control surfaces and actuators
• Type of main rocket engines (5 on the booster, 2 on the orbiter)
• Reaction control system
• Environmental control system for ascent and descent
• Duplex redundant GO2/GH2 auxiliary power units
• Avionics
• Tricycle landing gear with twin wheel units and deployment by gravity
• Diameter of rear nonintegral cylindrical LO2 and LH2 tanks (identical to Ariane 5 core stage)
While several subsystems on the booster may consequently be overdimensioned and not optimized with respect to mass, it was felt that the achievable overall cost savings would more than offset the incurred performance penalty, especially since the booster is much less sensitive to mass increases than the orbiter.
The booster is powered by five main engines and the orbiter by two main engines and two OMS engines, which together enable full abort capability in case of a single engine failure on either stage during all flight phases. For the subsonic cruise flight back to the launch site, the booster is also equipped with four airbreathing turbofan engines for single engine out capability, which are installed fixed with moveable inlets and exhaust nozzles in an upright position perpendicular to the flight direction in the front fuselage due to CoG considerations.
The orbiter incorporates an integral LO2 tank in the nose, followed by an integral LH2 tank, the payload bay and a nonintegral LO2 tank in front of the engine thrust structure.
The booster maintains the forward arrangement of the integral LO2 and LH2 tanks, but it accommodates another integral LO2 tank and a nonintegral LH2 tank instead of the payload bay and the rear nonintegral LO2 tank of the orbiter.
The main geometric of both winged stages data are summarized in Table 2. The fully reusable configuration is illustrated in Figure 10, and Figure 11 shows the semi-reusable version.
Table 2: Main Geometry Data of the FSSC-16 Fully Reusable Version
13 Stage Booster Orbiter Total Vehicle Length [ft] 132.2 132.2 Fuselage Length [ft] 131.2 131.2 Maximum Fuselage Width [ft] 21.0 21.0 Maximum Fuselage Height [ft] 21.6 21.6 Wing Span (including Winglets) [ft] 71.5 71.5 Total Wetted Area w/o Base [ft2] 13,548.5 13,548.5 Aerodynamic Reference Area [ft2] 4,365.8 4,365.8 Aspect Ratio 1.17 1.17 Fineness Ratio 6.15 6.15 Total Payload Bay Volume [ft3] N/A 6,179 3 Total LH2 Tank Volume [ft ] 18,787.4 11,332.5 3 Total LO2 Tank Volume [ft ] 7,228.9 5,177.1
The belly to belly configuration was maintained to enable unified structural interfaces on the booster for both versions and also corresponding structural hard points on both winged stages, since the divergent wing gap resulting from the dihedral of the lower surfaces of the wings alleviates the risk of aeroelastic effects during the combined ascent.
The identical shape of booster and orbiter reduces the number of aerodynamic configurations, which have to be qualified and controlled, from the typical value of three for a TSTO (booster, orbiter and compound) to two, with the one of the compound being aerodynamically symmetrical.
14 Figure 10: Layout of the FSSC-16 Fully Reusable Version (All Dimensions in mm)
15 Figure 11: Layout of the FSSC-16 Semi-Reusable Version (All Dimensions in mm)
16 Both winged vehicles have the following aerodynamic control surfaces:
• Central body flap for pitch control
• Left and right elevons for roll control as ailerons and additional pitch control as elevators
• Left and right vertical rudders on the winglets for yaw control
The CoG shift during ascent is alleviated in comparison to FSSC-9, since the booster mass is now considerably closer to that of the orbiter, and the relative mass change until staging is also smaller. The array of four separate main propellant tanks enables to perform propellant management for longitudinal CoG migration control during ascent.
The CoG of the empty vehicle for both stages is located at about 63 % of the body length, since the main differential mass impacts on the booster with respect to the CoG position, which result from the additional rocket engines in the rear and the turbofans in the front fuselage, compensate each other almost completely. The remaining difference can be eliminated by the respective integration of subsystems and components, such as the auxiliary power units.
Table 3 contains the top level mass budget for a mission into a near polar LEO, with (B) and (O) marking alternative positions for booster and orbiter, respectively.
Table 3: FSSC-16 Mass Budget for a Near Polar LEO Mission
Stage Booster Orbiter Tanks, Fuselage and Thrust Structure [lb] 29,613 24,525 Aerodynamic Surfaces [lb] 7,443 7,443 Propulsion System [lb] 42,943 15,267 Subsystems [lb] 12,112 12,093 Thermal Protection System [lb] 5,340 13,181 Design Margin [lb] 11,669 9,363 Ascent Propellant Mass [lb] 512,566 400,132 Flyback (B)/OMS (O)/RCS Propellant Mass [lb] 12,407 3,408 Propellant Residuals/Reserves/Losses [lb] 6,407 5,004 Upper Stage Mass (B)/ Payload Mass (O) [lb] 497,030 6,614 Launch Mass [lb] 1,137,530 497,030
17 The detailed dry mass breakdown for both stages, once again with (B) and (O) marking alternative positions for booster and orbiter, is compiled in Table 4.
Table 4: Dry Mass Breakdown of the FSSC-16 Booster and Orbiter
Stage Booster Orbiter Fuselage Structure [lb] 11,243 10,284
Front LO2 Tank [lb] 3,333 3,333
Front LH2 Tank [lb] 4,755 4,755
Rear LO2 Tank (B)/Payload Bay (O) [lb] 3,922 3,086
Rear LH2 Tank (B)/Rear LO2 Tank (O) [lb] 3,053 2,229 Interstage (B)/Payload Provisions (O) [lb] 1,323 2,756 Thrust Structure [lb] 1,984 838 Landing Gear [lb] 3,320 3,320 Wings and Winglets [lb] 6,967 6,967 Body Flap [lb] 476 476 Main Rocket Engines [lb] 23,336 9,334 Propellant Supply and Crossfeed [lb] 8,926 5,084 Airbreathing (B)/OMS Engines (O) [lb] 10,582 750 Reaction Control System Thrusters [lb] 99 99 Stage Separation System [lb] 2,529 666 Secondary Power System [lb] 1,642 730 Hydraulic System [lb] 791 791 Actuators [lb] 1,327 1,327 Environmental Control System [lb] 794 794 Avionics [lb] 329 329 Electrical System [lb] 1,380 1,380 Thermal Protection System and Hot Structures [lb] 5,340 13,181 Design Margin [lb] 11,669 9,363 Stage Dry Mass [lb] 109,120 81,872
18 5. Structures and TPS
The primary vehicle structure of the booster is composed of the following main elements:
• Nose
• Integral LO2 tank
• Intertank structure housing the forward RCS, the nose gear and the four turbofan engines
• Integral LH2 tank
• Nonintegral LO2 tank
• Nonintegral LH2 tank
• External shell
• Wings
• Winglets as vertical stabilizers
• Thrust structure
• Body flap
The turbofan engines are not present in the orbiter intertank structure, and instead of the rear nonintegral tanks on the booster, the orbiter includes the following elements:
• Payload bay
• Nonintegral LO2 tank
The nose, which experiences the highest thermal loads during reentry of the orbiter, is a monolithic hot structure made of C/C.
The cold primary structure of both stages is composed of Al frames and CFRP sandwich shells, with the nose section behind the hot structure manufactured from Inconel.
The cargo bay is made of CFRP sandwich walls stiffened by frames and is attached to the upper surface shell along the top edge.
All LO2 tanks are made from Al-Li 2195, and all LH2 tanks are made from CFRP. The forward integral tanks on booster and orbiter are identical to lower the development and
19 production cost through a design repeat. The integral LO2 tank has a volume of 2,009.4 3 3 ft , and the integral LH2 tank accommodates 11,332.5 ft .
3 The rear integral LO2 tank of the booster has a volume of 5,219.5 ft and the rear 3 nonintegral LH2 tank 7,454.9 ft , while the rear nonintegral LO2 tank of the orbiter has a capacity of 3,167.7 ft3.
The diameter of the rear nonintegral cylindrical LO2 and LH2 tanks on both winged vehicles is identical to that of the Ariane 5 core stage, so that common elements derived from this stage might be implemented on both stages in order to increase the commonality of elements between booster and orbiter and to make maximum use of Ariane 5 tooling and components for further cost reduction. The dry mass increase on the booster resulting from replacing the rear CFRP LH2 tank with one manufactured from Al-Li 2195 was assessed to be in the order of 1,000 lb, which might be well tolerable from a performance point of view.
All rear tank domes have a torospherical shape and are more rounded than the forward ones due to the additional hydrostatic loads they have to sustain.
All cryoinsulations are made of Rohacell foam, which is applied internally for the LH2 tanks and externally for the LO2 tanks.
The thrust structures on both stages are made of metal matrix composites. They are conical constructions, with the orbiter having a more compact cone than the booster.
Since the thrust level of the staged combustion main rocket engines was chosen identical to that of the Vulcain 2, the thrust structure of the booster could be used without modification in both variants.
The wing structure on booster and orbiter is identical to lower the development and production cost through a design repeat, even though the wing may consequently be overdimensioned for the orbiter, since it was adopted from an air launched stage.
The bodyflap, wing, aileron and winglet cold structures are made from Al ribs and spars with CFRP sandwich skins, and the wings have a carry through wing box, which is connected to corresponding fuselage frames. The wing and winglet leading edges are integral hot structures manufactured from C/C.
20 The stage attachment structure consists of two rear tripods and one front bipod. The thrust force is transferred from the booster to the orbiter via the tripods through compression loaded oblique struts. The struts are made from metal matrix composites.
The main interstage load paths are routed through the aft attachment struts, which are located close to the thrust structures and wing boxes of both stages. For an application of the booster in connection with an Ariane 5 core stage, this would necessitate a modification of the core stage.
The undersides of the fuselage front part as well as of the bodyflap, wings and winglets of the orbiter are covered by a C/SiC based shingle type TPS. Thermally less stressed surfaces are equipped with different varieties of a metallic multiwall TPS. For still lower thermally loaded areas, two versions of a lighter FEI type TPS are applied. The interfaces between hot and cold structures are protected by silica fiber felt layers.
The booster uses flexible external insulation and metallic multiwall TPS in thermally stressed areas. Due to the relative low staging conditions of the fully reusable system in comparison to the semi-reusable version, no dedicated TPS might be required at all for the booster in this application.
6. Propulsion
The main propulsion of FSSC-16 incorporates seven staged combustion rocket engines, five on the booster and two on the orbiter. The chosen staged combustion engines are less demanding with regard to performance parameters and smaller than the SSME or the Russian RD-0120, which reduces development risk and cost, and the technology level is comparable to that of the Japanese LE-7 engine.
Table 5 shows the main characteristics of these engines in comparison to the Vulcain 2.
The thrust vectors of the main engines are aligned in parallel in each stage in nominal operation, rather than aimed through the CoG, in order to maximize the performance of the propulsion system. The five booster engines are arranged in an X-shaped cluster. To provide single engine out capability, the orbiter has two engines. In contrast to the original arrangement on the HORUS orbiter shown in Figure 8, the two orbiter engines are arranged above each other instead of side by side, in order to avoid asymmetric thrust in case of a single engine failure or shutdown for maintaining the acceleration limit towards the end of the powered orbiter ascent.
21 Table 5: Main FSSC-16 Rocket Engine Data in Comparison to the Vulcain 2
Engine Vulcain 2 FSSC-16 Engine Cycle Gas Generator Staged Combustion Mixture Ratio 6.1 6.6 Mass Flow Rate [lb/s] 703.0 661.4 Chamber Pressure [psia] 1,668 2,176 Nozzle Area Ratio 58.5 65.84 Sea Level Specific Impulse [s] 327 346.7 Vacuum Specific Impulse [s] 434 447.1 Sea Level Thrust [lbf] 229,000 229,224 Vacuum Thrust [lbf] 303,384 295,609 Mass [lb] 4,266 4,500 Length [ft] 11.2 11.7 Diameter [ft] 6.7 7.1
The nominal liftoff acceleration is 1.4 g, while the residual value in case of a single engine failure is 1.2 g to assure safe takeoff and controllability for an abort. A single engine failure is compensated by engine gimbaling, with the remaining six engines ensuring the minimum liftoff acceleration at their nominal power level. The maximum axial acceleration limit during powered ascent of 3.5 g can be maintained by throttling both or switching off one orbiter engine after stage separation, since the limit is not reached before staging.
The throttling range down to 26%, which is needed for the central booster rocket engine prior to staging in order to achieve force equilibrium between booster and orbiter for safe separation conditions even in case of a single engine failure on the orbiter, can be accomplished by understochiometric operation with fuel rich combustion.
The orbiter engines are fuelled from the booster tanks during the combined ascent via crossfeed. The tanks in both stages are arranged in such a way that the transfer of LO2 as the denser liquid from the booster to the orbiter tanks is induced by gravity respectively the vehicle acceleration, since both LO2 tanks of the booster are in flight direction located above that of the orbiter. Active pumping is only required for the significantly less dense LH2, thus minimizing the necessary pump power.
22 The crossfeeding is performed from the booster tanks into the orbiter tanks, so that the orbiter engines are always fed from their tanks, and no switching device, which could lead to waterhammer effects during cutoff, is required in the orbiter manifold. The constant replenishment of the orbiter tanks also prevents propellant stratification, especially of the LH2.
The LO2 and LH2 crossfeed installations on both stages are located next to each other in the rear fuselage beneath the payload bay respectively the nonintegral tanks. This position was chosen to remove the connections as far as possible from the front separation pyrotechnics to reduce the risk of ignition of propellant residuals.
The pressurization of the LH2 tanks is done with GH2, and the pressurization of the LO2 tanks is accomplished by using liquid Helium, which is evaporated by a heat exchanger before being fed into the LO2 tanks. This heat exchanger is integrated into the LO2 turbopump turbine exhaust pipe. The spherical Helium tank is embedded in one of the
LH2 tanks to save thermal insulation. The GH2 is taken from the thrust chamber cooling loop of the main engines. Due to the relatively low pressure in the main propellant tanks, boost pumps with high anti-cavitation characteristics and coaxial shaft construction are foreseen. In order to achieve redundancy and avoid the potential for single point failures of a central tank mounted boost pump, these boost pumps are mounted on all engines and located directly before the main pumps in a cascade arrangement. The high power level of these pumps requires hydraulic turbines, which are driven with fluids taken from the main LH2 and LO2 pump outlets.
The orbiter is equipped with two OMS engines to enable single engine out capability, which are assumed to be HM-7B derivatives modified for reusability. The primary engine data are compiled in Table 6.
The OMS engines are installed right and left of the linear main engine array in a horizontal plane on the main thrust structure with their nominal thrust vectors pointing through the vehicle CoG on orbit. This arrangement eliminates the potential induction of moments resulting from an OMS engine failure and avoids interference with the main engine gimbaling and body flap deflection envelopes. The engines can be gimbaled to account for CoG shifts due to payload deployment.
Table 6: HM-7B Derivative OMS Engine Data
Cycle Gas Generator
23 Mixture Ratio 4.56 Chamber Pressure [psia] 520.7 Vacuum Specific Impulse [s] 447.9 Vacuum Thrust [lbf] 14,135 Mass [lb] 342 Length [ft] 6.6 Diameter [ft] 3.2
The OMS engines may be used to support the two orbiter main engines during stage separation to achieve zero relative acceleration between both stages, in order to minimize the performance losses due to the staging procedure and the associated throttling of the central booster engine. The OMS engines may also be activated in case of an abort maneuver to support the orbiter main engines by increasing the available thrust as well as reducing the onboard propellant load, as was done in case of the Space Shuttle during an abort to orbit on the Spacelab 2 mission in 198510. The OMS acceleration is 0.24 g in comparison to 0.05 g on the Space Shuttle Orbiter.
The booster is equipped with four Hydrogen fuelled turbofans for the powered subsonic flyback, which correspond to the Eurojet EJ200 used on the Eurofighter EF2000 regarding size, performance and general design. A number of four engines instead of the two on the FSSC-9 booster was chosen, since this allows to achieve single engine out capability for the airbreathing cruise phase with less installed excess thrust and associated engine mass and enables a better accommodation due to the smaller individual engine size.
The adaptation of conventional turbojet and turbofan engines for utilizing Hydrogen has been repeatedly successfully demonstrated in flight tests in the USA in 1956 with a converted B-57 and in Russia in 1988 with the TU-155 demonstrator11. The introduction of Hydrogen as fuel for the airbreathers reduces the specific fuel consumption and therefore the required propellant mass by a factor of about 2.77 in comparison to Kerosene and thus minimizes the influence of the flyback propellant mass as a driver for the takeoff mass of the vehicle. It requires only minor modifications to the combustion chambers and fuel system components like fuel pumps, fuel pipes, control valves and injection systems of current turbojets or turbofans, thus limiting the technology
24 development risk. Since the use of Hydrogen enables a unified propellant management system as well as simplified operations for the booster and is also expected to improve the in flight engine start characteristics, it is regarded as preferable over Kerosene.
Potentially critical points of the design solution for the integration and activation of the airbreathers in the FSSC-9 booster by swinging out nacelles on the fuselage aft underside were the mechanical complexity and the structural interface of the deployment mechanism, the location of the fuel tanks and the necessity to penetrate the TPS with a large cover. Therefore a fixed engine installation with only deployable air inlets and exhaust ducts was considered preferable.
The chosen arrangement is a rigidly mounted engine installation in an upright orientation
(perpendicular to the direction of flight) between the forward integral LO2 and LH2 tanks in the front fuselage right and left of the nose landing gear well. The accommodation in pairs on either side corresponds to two engine bays of the Eurofighter EF2000. The resulting thrust asymmetry in case of a single engine failure is reduced by the location of the engines close to the longitudinal vehicle axis. The forward location was chosen to balance the mass of the additional three rocket engines in the back of the vehicle in comparison to the orbiter and thus be able to maintain the CoG position necessary for the aerodynamic flyback. The motivation for the internal accommodation of the engines was the possibility to have one common aeroshape for both stages during hyper- and supersonic flight, which would have been precluded by having the airbreathing engines in external pods. The fixed installation of the engines was selected in order to reduce the design complexity by eliminating the need for heavy mechanisms for engine deployment and retraction, as well as to prevent having large openings in the external structure and simplify the Hydrogen supply by eliminating the demand for flexible feedlines. The location of the inlets on top of the fuselage also minimizes the risk of foreign object damage/ingestion during landing. The engine integration is shown schematically in Figure 12.
25 Figure 12: Airbreathing Engine Integration in the Booster
Although the installation is unconventional, it is notably less radical than the engine installation envisaged for the so-called RIVET supersonic STOVL fighter concept proposed by Lockheed12. This design incorporated a backwards installed engine, which was fed by an inlet diffuser that turned the airflow by 180°, while the movable exhaust nozzles again reversed the flow by 180° in cruise flight and could be swiveled down for vertical flight and even slightly forward for vectoring in forward flight. The RIVET configuration is shown in Figure 13.
Figure 13: RIVET Configuration and Engine Accommodation
26 While the reversal of the exhaust flow in the RIVET layout is analogous to the flow path in the rotatable nozzles of the Pegasus turbofan engine used in the Harrier and was not expected to cause complications, the inlet bend was perceived as a potentially significant problem and was studied in more detail by a CFD analysis as well as tests with a duct model. Internal turning vanes were needed to keep the flow attached, but the flow turning could be accomplished with acceptable front face distortion. Both the inlet and the nozzle were evaluated to lead to about 3-5% performance loss each. The resulting maximum thrust loss of 10% for RIVET is consistent with an estimate of 5% loss of the overall propulsive efficiency assessed for the engine installation on the FSSC-16 booster, since both the inlet and the nozzle flow redirection on the booster is only half that of the respective values of RIVET. The findings for RIVET can therefore be considered a validation of the chosen approach.
During the ascent and the unpowered first portion of the booster return, the air inlets and the engine exhausts are retracted respectively closed by moveable ramps. A potential design for the inlet is a deployable Pitot airscoop similar to the construction used on the Tomahawk BGM/UGM-109 Cruise Missile, which is at launch retracted into the fuselage flush with the outer surface.
The afterburners and variable exit nozzles of the military turbofan engines are replaced with moveable nozzles with a fixed exit area. A local reinforcement of the TPS on the booster underside downstream of the nozzles may be required due to the tangential flow of the exhaust to the fuselage surface.
The inlets and nozzles are opened during the glide down to cruising altitude, and the engines are started using internal electrical power, supported by dynamic pressure, which initiates a windmilling of the turbofans once air flows through the engine. The peak power required for the in flight engine startup on the order of 100 kW per engine is provided by auxiliary power units in conjunction with backup batteries.
The Hydrogen has to be supplied to the engines in gaseous form, so that a heat exchanger respectively vaporizer is required in the propellant supply and conditioning system. It might however also be feasible to utilize already evaporated fuel present in the tanks, thus taking advantage of the residuals.
The spin of the rotating parts of the turbofans may lead to flight mechanical interactions during maneuvers, which differ from the effects on aircraft with conventional engine
27 installations due to the orientation of the axis of rotation perpendicular to the direction of flight and would have to be accounted for by the flight control system. Since no major maneuvers are expected during the powered flight, which will essentially be performed on a straight flight path, these effects will however most probably be negligible.
The upward pitching moment caused by the offset of the vectors of the impulse of the incoming air and the impulse of the exhaust gases can be trimmed with the bodyflap.
The acceleration loads on the engine normal to the longitudinal axis in combination with strong vibrations during ascent are compatible with the load orientation encountered in military aircraft when flying tight maneuvers resulting in high normal g loads. The engine fluid circuits will however probably have to be adapted to the sustained operation in an upright position.
Both stages are equipped with an RCS, which is based on GO2/GH2 fuelled hot gas thrusters. The propellants are stored in spherical high pressure tanks. Two separate RCS engine supply systems, which each feed two thruster clusters, are located in the upper nose compartment in front respectively in the rear of both stages to decrease the pressure losses associated with long feed lines. Each RCS module contains 14 thrusters, which are arranged in two redundant clusters of 7 thrusters each. The general RCS layout and the RCS thruster positions on both stages are identical. Each one of the front clusters has one thruster oriented towards the front, one left, one right, two up and two down. The back clusters have each one thruster oriented towards the back, one left, one right, two up and two down. Having the same number of thrusters in each of the four clusters enables the use of identical propellant manifold valve sets on both stages.
7. Subsystems
All subsystems were dimensioned for a mission duration of 16 orbits, which corresponds to about 24 hours.
The hydraulic system incorporates electrohydraulic actuators with centralized power supply and independent individual electric pumps for operating the flight controls, the engine thrust vectoring and the nosewheel steering. Electromechanical actuators are used for the landing gear control and mechanical elements, which are operated in orbit, such as the orbiter payload bay doors. All actuators are redundant. The landing gear is equipped with electrical brakes.
28 The electrical system includes generators, control units, DC/DC converters, power control relays and harnesses. For the secondary power system of the orbiter, a duplex redundant system based on three redundant GO2/GH2 fuelled gas generator driven turbine auxiliary power units with generators including self supplied GO2/GH2 conditioning was selected for supplying hydraulic power during ascent and descent. The electrical power for the orbiter is provided by a redundant layout based on three fuel cells during all mission phases. These fuel cells are replaced by backup Silver-Zinc batteries, which supplement the auxiliary power units during ascent, respectively turbofan mounted and driven generators for the booster. Depending on the peak power demand for turbofan startup, the auxiliary power units on the booster could be eliminated, if the batteries alone were found to be sufficient.
Thermal conditioning of the internal vehicle structure and subsystems is performed by a two loop regenerative thermal management system using LH2 or GH2 in the primary loop and water in the secondary loop. The waste heat generated by the hydraulics, the main propulsion, the payload, the avionics and the fuel cells is rejected via heat exchangers.
Due to cost, mass and flexibility considerations, the data management system including the guidance and navigation system, the flight control system and the health management system was conceived as a distributed architecture and designed for a high degree of autonomy.
The landing gear is a conventional tricycle type undercarriage. Both the main and nose gears are fitted with twin wheel units. The gear is retracted using external force on the ground and is deployed by gravity. The landing gear mass is identical for both stages, which means that it is slightly overdimensioned for the orbiter.
The stage separation system consists of four solid separation rocket motors, which are located in the front and rear fuselage of the booster. The forward motors are located parallel and adjacent to the airbreathing engines between the two forward integral tanks, while the rear motors are installed in the main engine bay behind the wing. The location and orientation of the thrust vectors was selected with respect to the need to avoid nearfield plume impingement on either the orbiter or the Ariane 5 core. The stage attachment struts are severed at separation by pyrotechnic shear bolt devices on the orbiter. Similar devices are foreseen on the booster in case a retractable and partly reusable attachment system is not feasible. An active closure system is baselined for the corresponding openings.
29 8. Mission Profile and Performance
The baseline mission profile of the fully reusable system consists of:
• Combined vehicle ascent powered by the main engines of both stages until separation
• Stage separation
• Unpowered curve flight of the booster and heading alignment for return to the launch site
• Powered return cruise of the booster
• Final approach and landing of the booster
• Continued ascent of the orbiter until main engine cutoff
• Coast phase in transfer orbit
• Circularization in target orbit via OMS
• Payload checkout and deployment
• Deorbit maneuver via OMS
• Unpowered descent in transfer orbit
• Atmospheric reentry and return glide
• Final approach and landing of the orbiter
The qualitative flight segments for the booster in the semi-reusable version are basically identical to those of the fully reusable system, while the mission profile of the expendable core corresponds to that of the Ariane 5.
Trajectory simulations were performed for both design missions of the fully reusable TSTO as well as for a LEO and a GTO mission of the semi-reusable version.
The AoA is kept zero for the combined ascent phase to minimize the ground track and the flyback distance the booster has to cover during the return flight. Consequently, pure drag was assumed for the ascent aerodynamics during the combined flight phase. Ascent control is performed by engine gimbaling in connection with the aerodynamic control surfaces on both stages. After stage separation, both stages utilize aerodynamic lift to optimize their performance.
30 The longitudinal acceleration is limited to 3.5 g. Staging of the fully reusable system occurs 111 seconds into the flight at Mach 4 in 17.8 NM altitude for the near equatorial LEO and at Mach 4.2 in 18.4 NM altitude for the near polar LEO.
Staging is performed at Zero differential acceleration between both stages and an AoA close to Zero. The acceleration equilibrium is achieved by shutting down four of the booster engines and throttling the remaining central one. For the nominal stage separation with the two orbiter OMS engines working parallel with the two main engines, a throttling degree of 50% is required, while for a single engine out on the orbiter this value has to be decreased to 26%. This throttle range allows to follow the same separation logic on the semi-reusable version, where a power setting of 34% is required.
In case of an abort due to an engine failure, the staging conditions can be raised to a higher altitude in order to assure safe separation by prematurely terminating the crossfeed between both stages and running the orbiter engines independently on internal propellants, so that the booster burn time is prolonged. This strategy also allows to achieve an acceleration equilibrium between both stages and safe separation even in case of an orbiter engine failure by partially depleting the orbiter tanks before staging, thereby increasing the stage thrust/weight ratio.
The potential abort alternatives include return to launch site for the booster and return to launch site, abort downrange, abort once around and abort to orbit for the orbiter. In order to avoid off-site landings necessitating an expensive ferry back to the launch site, the last three alternatives are preferred for the orbiter. The overall dimensions and dry masses of both stages allow however to transport them on top of a 747 Jumbo Jet analogous to the Shuttle Carrier Aircraft, so that the orbiter could be retrieved in case of a downrange abort.
Flight time until orbiter main engine cutoff is 425 seconds, when the vehicle enters a 54 NM ∗ 135 NM transfer ellipse. Both the near polar and the near equatorial target orbit are achieved about half an hour after liftoff with insertion by OMS. Typical payloads require about 6 hours, corresponding to 4 orbits, for activation, checkout, establishment of datalink and deployment. For the low inclination near equatorial orbit, the orbiter can return to the launch site from orbit during every revolution. For the high inclination near polar orbit, the ground track has to be phased with the Earth’s rotation, so that the return to Kourou is only possible from the 7th respectively the 8th and again from the 15th
31 respectively the 16th revolution. This corresponds to an on orbit mission duration of 12 hours for the first two consecutive return possibilities and 24 hours for the second two. The subsystems were consequently dimensioned for a maximum exoatmospheric mission duration of one day.
The semi-reusable version stages 155 seconds into the flight at Mach 6.3 in 29.9 NM altitude for a LEO mission and at Mach 6.5 in 30.3 NM altitude for a GTO mission.
For the semi-reusable version, the only abort alternative is a return of the booster to the launch site, but it has no complete abort capability coverage for the booster during the combined ascent in case of a failure of the single core stage engine. For full abort capability in case of a single engine failure on either stage during the entire ascent the Ariane 5 core stage would have to be reconfigured to a twin engine configuration and propellant crossfeed between booster, and core stage would have to be implemented, which was found to be not cost effective.
A reliability analysis assessed the overall risk of loss of vehicle for the fully reusable system to be 0.2%, while the risk of loss of vehicle for the semi-reusable version was found to be over 1% due to the relatively low reliability of the expendable components.
The booster descent and return was studied in detail for the GTO mission of the semi- reusable version, which constitutes the most demanding case with the highest staging conditions dimensioning the required flyback propellant loading capability.
After separation, the booster flies an unpowered curve to initiate the return to the launch site. The normal acceleration during the curve flight for heading alignment was limited to 2.5 g. The resulting maximum cruise flight distance for the ascent into a GTO orbit is 296 NM. The curve and glideback after staging until reaching the cruise conditions lasts about 12 minutes, and the subsequent powered flyback phase including landing lasts about one hour. The total maximum flight time of the booster from liftoff until touchdown is around 1 hour 15 minutes.
The propelled booster flyback is performed in 20,000 feet altitude at Mach 0.5. The flyback conditions were chosen to allow flight above weather formations in order to minimize the risk of icing. In case of a failure of one of the airbreathing engines, the altitude will have to be reduced to 10,000 feet in order to increase the thrust available from the remaining engines to the required level. The nominal cruise is performed at an AoA of about 8°, which corresponds to a trimmed L/D of about 5.
32 Opposed to the orbiter, the booster landing is performed powered. Since the thrust vector partly supports the weight during the final flare before touchdown, the vertical landing speed is reduced in comparison to the orbiter. This effect counters the higher landing mass of the booster, and since the orbiter has also to be able to land with the full payload on board in an abort case, it allows to use the same landing gear on both stages without a significant mismatch. Landing speeds below 330 feet/second can be achieved with an AoA of above 10° for the booster, while on the orbiter, an AoA of above 9° is sufficient due to the lower planform loading.
Additional simulations showed, that due to the low staging conditions of the fully reusable system no propelled cruise flight would be required at all for the booster in this application, if the structures were dimensioned to withstand normal accelerations in the order of 4 g during the return curve.
The fully reusable system exceeded the FESTIP mission requirements, enabling a payload capability of 6,614 lb into a near polar LEO for a launch mass of 1,137,530 lb and 22,046 lb into a near equatorial LEO for a launch mass of 1,153,000 lb The semi- reusable system enabled slight performance increases for the Ariane 5 in both LEO and GTO.
A sensitivity analysis of the fully reusable design performed for a constant dry mass ratio showed the orbiter to be close to the theoretical optimum for the given booster size, while the booster size, which was dictated by the compatibility requirement with Ariane 5, is not optimized for the fully reusable application with respect to the achievable payload ratio. It has however to be kept in mind, that the actual cost/performance optimum is driven by the commonality between both stages, which more than compensates deviations from the theoretical optimum value with respect to mass ratios.
9. Ground Infrastructure and Operations
In order to enable a high degree of flexibility and turn around efficiency as well as resiliency and redundancy, a parallel turn around scheme allowing the simultaneous processing and preparation of two launchers was defined for both the fully and semi- reusable versions.
Launch and landing site for all versions is Centre Spatial Guyanais in Kourou, French Guyana. This location made closed air conditioned buildings for vehicle handling mandatory. For the semi-reusable version, the current integration and preparation
33 methods of the Ariane 5 core were maintained, while for the fully reusable stages the philosophy of aircraft like ground operations was adopted as far as possible.
After landing on the runway under supervision of a Control Center the booster is either brought to a Booster Preparation Building in case of the fully reusable system respectively to a Launcher Assembly and Preparation Building in case of the semi- reusable version for on line maintenance and preparation or the Orbiter/Booster Overhaul Building for off line refurbishment.
In the Launcher Assembly and Preparation Building the Ariane 5 core is assembled and prepared and the booster is maintained, erected and mated with the expendable core. The compound is then transferred on a mobile Launch Table into the Final Integration Building. The payloads are prepared in the Payload Preparation Building, and after transfer in a cleanroom container to the Final Integration Building they are integrated into the launcher.
In case of the fully reusable system, analogous to the booster after landing the orbiter is either brought to the Orbiter Preparation Building for on line maintenance and preparation or the Orbiter/Booster Overhaul Building for off line refurbishment. Additional stages can be parked in an Orbiter/Booster Hangar.
The payload is integrated horizontally into the orbiter in the Orbiter Preparation Building by means of a cleanroom container, which is used to transfer the payload from the Payload Preparation Building and is docked to the payload bay for payload integration.
After the separate parallel processing, both stages are erected, placed on the mobile Launch Table and mated in the Uprighting and Assembly Building.
Following final checkout in the Final Integration Building in case of the semi-reusable version respectively in the Uprighting and Assembly Building in case of the fully reusable system the launcher is transferred on the mobile Launch Table to the Launch Pad and fuelled from a stationary Umbilical Building.
All engines are ignited on the ground, and the booster is held down by clamps, until proper engine function in all stages is verified and the nominal liftoff thrust has built up. For the reusable orbiter as well as the Ariane 5 core, the thrust/weight ratio of the individual stage on the pad is lower than one, so that no holddown mechanism is required for these stages.
34 10. Cost and Programmatics
Development, production and operations cost assessments were performed for the fully and semi-reusable versions of FSSC-16, but the absolute cost figures for all FESTIP concepts have been classified as confidential information by ESA and can therefore not be disclosed in this report. In comparison to FSSC-9, the development cost of the fully reusable system was however reduced by more than 30% through the high level of commonality between booster and orbiter on the component level, and the cost per flight was reduced by 17% relative to that of FSSC-9, while at the same time the performance in both target orbits was improved. The development cost of FSSC-16 was the lowest of all fully reusable FESTIP concepts comprising various SSTO and TSTO designs, which would enable later payload return missions and crewed orbital operations.
The following technologies were identified as critical for the feasibility of the concept:
• Cold CFRP primary structures
• Integral and nonintegral CFRP LH2 tanks
• Integral and nonintegral Al-Li LO2 tanks
• C/C and C/SiC hot structures
• C/SiC shingles, FEI and metallic multiwall TPS
• Reusable cryoinsulation
• Reusable staged combustion rocket engines with 2176 psia chamber pressure
• Hydrogen fuelled derivatives of current turbofans with upright installation
• Cryogenic propellant crossfeed
• Control of combined ascent
• Stage separation
• Nondestructive inspection and health monitoring technology
The basic concept of FSSC-16 allows a gradual evolution, starting with a reusable booster for a semi-reusable application and progressing towards a reusable orbiter for a fully reusable system, while at the same time minimizing the development effort by achieving a high degree of commonality between both reusable stages. The gradual development and introduction of reusability into the system, starting with the booster, is
35 motivated by the less demanding flight environment of this stage together with the high number of installed engines and the associated investments as opposed to the Ariane 5 core or the orbiter, which enables higher savings by introducing reusability. The booster allows the incremental improvement of technologies yielding maximum synergy with a future orbiter.
Due to the lack of a second design loop, the booster propellant tanks are oversized by about 10%. This would enable the following alternatives for further design evolution:
• Introduction of a circular, nonintegral rear LO2 tank with lower development risk and possibly higher structural commonality between booster and orbiter
• Shortening of the tanks to create alternatives for the turbofan installation
• Separation of the rear nonintegral LH2 tank and introduction of a dedicated flyback propellant tank to prevent propellant sloshing, if necessary
• Increase of the usable ascent propellant mass by about 48,500 lb for performance enhancement
• Elimination of the airbreathers and use of some of the additional propellant to initiate a ballistic flyback by a rocket impulse ‘tossback’ maneuver13 of the booster after staging
Possible improvements of the basic layout of both winged stages, which could also not explored in the initial design loop, concern the introduction of a simplified fuselage with completely rotationally symmetrical tanks and of a single vertical tail/stabilizer instead of the winglets.
36 In addition, a modular Ariane 5 derived HLLV as proposed by CNES for lunar missions14 could be created by replacing the originally foreseen four solid boosters by two reusable winged boosters, analogous to the arrangement shown in Figure 7, in combination with a H-620 cryogenic core stage and a H-80 cryogenic upper stage. The resulting configuration is shown in Figure 14. The engine bay for five Vulcain engines needed for the core stage could be derived from the corresponding component of the booster.
Figure 14: Ariane 5 Derived HLLV with Twin Reusable Boosters
37 11. Acronyms
ADV Ariane Derived Vehicle
Al Aluminum
Al-Li Aluminum-Lithium
AoA Angle of Attack
C/C Carbon/Carbon
C/SiC Carbon/Silicon Carbide
CFD Computational Fluid Dynamics
CFRP Carbon Fiber Reinforced Plastic
CoG Center of Gravity
DASA DaimlerChrysler Aerospace
DC/DC Direct Current/Direct Current
EARL European Advanced Rocket Launcher
FEI Flexible External Insulation
FESTIP Future European Space Transportation Investigations Programme
FSSC FESTIP System Study Concept
GH2 Gaseous Hydrogen
GO2 Gaseous Oxygen
GTO Geostationary Transfer Orbit
HTHL Horizontal Takeoff Horizontal Landing
IMI Internal Multilayer Insulation
L/D Lift/Drag
LEO Low Earth Orbit
LH2 Liquid Hydrogen
LO2 Liquid Oxygen
38 NASA National Aeronautics and Space Administration
OMS Orbital Maneuvering System
RCS Reaction Control System
RIVET Reverse Installation Vectored Engine Thrust
RLV Reusable Launch Vehicle
RRL Reusable Rocket Launcher
RVD Rendezvous and Docking
SSME Space Shuttle Main Engine
SSTO Single Stage to Orbit
STOVL Short Takeoff Vertical Landing
TPS Thermal Protection System
TSTO Two Stages to Orbit
USA United States of America
VTHL Vertical Takeoff Horizontal Landing
WLC Winged Launcher Configurations
12. Literature 1 R. Stadler: Results for a Fully Reusable TSTO-Launch Vehicle Concept, AIAA-98-1504
39
2 W. Westphal, K.-W. Kalk, G. Greger: Views on the Evolution of European Space Transport: The Reference Concept EARL II, Proceedings of the Second European Aerospace Conference on Progress in Space Transportation, May 22-24, 1989, Bonn- Bad Godesberg, Germany, ESA SP-293, August 1989, pp. 391-396 3 R. Reichert: Future Reusable Space Launcher Systems: A European View, Proceedings of the First International Conference on Hypersonic Flight in the 21st Century, September 20-23, 1988, University of North Dakota, Center for Aerospace Sciences, Department for Space Studies, Grand Forks, ND, USA, pp. 437-444 4 J. Martin: Two-Stage Earth-to-Orbit Vehicles with Series and Parallel Burn, AIAA-86- 1413 5 H. Lacaze, J. Bombled: Recoverable Rocket Launcher Configuration Selection, IAF 90-177 6 P. Terrenoire, B. Massé, G. Präger: An Overview of the Reusable Rocket Launchers Based on Near Term Technologies, IAF-94-V.3.539 7 R. Parkinson: Cost Sensitivity as a Selection Issue for Future Economic Space Transportation Systems, AIAA-95-6122 8 H. Kuczera, H. Hauck: The German Hypersonics Technology Programme, Status Report 1992, IAF-92-0867 9 W. Berry, H. Grallert: Performance and Technical Feasibility Comparison of Reusable Launch Systems: A Synthesis of the ESA Winged Launcher Studies, IAF-95-V.3.03 10 Abort-to-Orbit Incident Will Intensify Shuttle Engine Procedure Reviews, Aviation Week & Space Technology, August 5, 1985 11 H. W. Pohl: Hydrogen and other Alternative Fuels for Air and Ground Transportation, John Wiley & Sons, 1995 12 D. P. Raymer, Y. T. Chin, Y. F. Kiefer, K. J. Hajic, J. L. Benson: Supersonic STOVL: The future is now, Aerospace America, August 1990 13 L. E. McKinney: Vehicle Sizing and Trajectory Optimization for a Reusable “Tossback” Booster, MDC H1588, June 1986 14 P. Jorant: Ariane 5 Family, AIAA 93-4131
40