Third Way of Development of Single-Stage-to-Orbit Propulsion

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1 JOURNAL OF PROPULSION ANO POWER. Vol. 16, No. 1, January-February 2000 Third Way of Development of Single-Stage-to-Orbit Propulsion V. V. Balepin* MSE Technology Applications, Inc., Butte, Montana M. Maita National Aerospace Laboratory, Tokyo 182, Japan and S. N. B. Murthy Purdue University, West Lafayette, Indiana Considering a number of options for a space-launch vehicle-propulsion system, between advanced rocket motors and airbreathers, in particular, thermally integrated rocket-based combined-cycles, a new cycle that has been given the name KLIN (meaning wedge in Russian), is proposed as the "third way." It consists of a combination of a liquid rocket engine and a deep-cooled turbojet (with oxygen addition to atmospheric air). Currently, the KLIN concept is considered for application with a vertical takeoff horizontal landing vehicle, with the turbojet and the rocket engine optimized individually, with a possibility for incorporating an aerospike-type nozzle. Retaining a rocket trajectory up to Mach 3, the turbojet and rocket engine are assumed to operate together from takeoff with a gradual reduction in the deep-cooled turbojet output, finally terminating the turbojet at Mach 6. It can be shown that the KLIN can be manufactured with available or foreseeable technology, and provides a combination of engine weight and specific impulse that yields twice the payload mass fraction, in addition to other advantages, compared with a rocket-engine-operated vehicle. INTRODUCTION IT is universally agreed that the opportunities in space and space technologies mar be realized in full only if the space-launch vehicle becomes recoverable at any stage of the mission, and with only a modest amount of refurbishing, is reusable over a sufficiently large number of mission flights. It is also agreed that propulsion is the major aspect of the vehicle that must be improved, whether the goals are long-range or nearterm. For the long-term goals, there are sufficient arguments to show that the ultimate objective could be a horizontal takeoff single-stage-to-orbit (HTO SSTO) or TSTO system that is recoverable, except for the fuel consumed, and including the payload, unless it has been delivered; and that the largest part of the trajectory should be traversed using ambient atmospheric air, whereas rocket motors that are used should have the most effective specific thrust and specific impulse. For the near term, on the other hand, there is considerably greater debate in regard to propulsion and takeoff and landing. The parameters that can lead to success are, again, a major increase in specific impulse and thrust and a substantial decrease in the dry weight fraction of the vehicle. Two approaches have resulted from the near-term-use studies: 1) obtaining the best rocket motors, and 2) developing rocket-based combined-cycle engines 1. Under approach 1, a variety of options have been considered, such as fuel improvement, dual fuel, dual mixture ratio, dual position nozzle, dual throat, etc. However, there is a concern, e.g., Refs. 2 and 3, that the mixed-mode rocket propulsion mar only be marginally better than conventional, high-pressure hydrogen-oxygen rockets. However, more definitive positions may evolve from the X-33 project in the U.S. and somewhat similar studies elsewhere. Regarding approach 2, there is considerable scope for innovative improvements, as shown, e.g., in Refs. 4 and 5, although many of the required technologies must proved in themselves and as part of an integrated propulsion system for a chosen mission objective. Three choices of airbreathers for combining with VTO rockets are discussed in Refs. 5-7; they are based on the optimal use of hydrogen properties: 1) a liquid air cycle engine with mixed (air + oxygen) oxidizer (LACE-M), 2) LRE combined with an air-separation system, and 3) a deep-cooled turbojet thermally integrated with LRE. The hydrogen usage in each case brings about thermal integration of the LRE and the airbreather. With this background, one can ask if there is an option, a "third way" in the near term, between the airbreathing SSTO and the classic rocket: in particular, if there is a combined-cycle propulsion that can be realized in the near term without a need for large leaps in technology. The objective of this paper is to present such an option in the form of a combined cycle that is designated KLIN (meaning wedge in Russian). [KLIN is a trademark of V. Balepin and MSE Technology Applications, Inc., for a combined deep-cooled turbojet-liquid rocket engine (DCTJ-LRE) propulsion]. 1 Presented as Paper at the AIAA 7th International Space Planes and Hypersonic Systems and Technologies Conference. Norfolk, VA, November 1996; received 6 February 1998; revision received 30 November 1998; accepted for publication 5 December by the American Institute of Aeronautics and Astronautics, Inc. AlI rights reserved. *Staff Scientist. Senior Member AIAA. Head of Spaceplane System Studies. Senior Member AIM. Senior Reseacher. Associate Fellow AIAA.

2 References 1 Escher, W. J. D., "A Winning Combination for Tomorrow's Spaceliners, Aerospace America, Feb. 1996, pp Mansci, D., and Martin, J. A., "Evaluation of Innovative Rocket Engines for Single-Stage Earth-to-Orbit Vehicles," Journal of Propulsion and Power, Vol. 7, No. 6, 1991, pp Mansci, D., and Fina, A., "Advanced Rocket Propulsion Systems for Reusable Ballistic Single-Stage-to- Orbit Vehicles BETA and Delta-Clipper, AIAA Paper , June Czysz, P. A., "Rocket Based Combined Cycle (RBCC) Propulsion Systems Offer Additional Options, 11th International Symposium on Air Breathing Engines (ISABE), Tokyo, Sept Balepin, V. V., Yoshida, M., and Kamijo, K., "Rocket Based Combined Cycles for Single Stage Rocket, SAE Aerospace Atlantic Conference, Dayton, OH, Society of Automotive Engineers, Paper , April Balepin, V. V., Folomeev, E. A., Galkin, S. M., and Tjurikov, E. V., "Rocket Based Combined Cycles for Vertical Take-Off Space Vehicles, AIAA Paper , Apri Balepin, V. V., Maita, M., Tanatsugu, N., and Murthy, S. N. B., "Deep Cooled Turbojet Augmented with Oxygen (Cryojet) for a SSTO Launch Vehicle, AIAA Paper , July Baumgarther, R.I., and Elvin, J. D., "Lifting Body--An Innovative RLV Concept," AIAA Paper , July Balepin, V. V., and Maita, M., "KLIN Cycle: Combined Propulsion for Vertical Take Off Launcher, AlAA Paper , July Tanatsugu, N., Balepin, V., Sato, T., Mizutani, T., Hamabe, K., and Tomike, J., "ATREX Engine Development: First Practical Experience of Precooled Turbomachinery, Proceedings (1 the 5th International Symposium Propulsion in Space Transportation, Association Aeronautique et Astronautique de France, France, 1996, pp

3 Potential Cost Reduction by Ultra-Low-Cost Expendable Launch Vehicle Technology R. L. Sackheimt and P. Dergarabedian 2 Abstract The high cost of access Lo space begins with the current availability of very expensive ($3,000 - $5,000/lb payload to LEO) launch systems. This paper discusses and compares two distinct alternatives for making significant (nearly order-of-magnitude) reductions in launch costs. An additional benefit to dramatic reductions in the cost of space transportation would also be significant cost reductions of the payloads (i.e., spacecraft) by designing for minimum cost and not for minimum weight and maximum performance. One approach proposes a reusable single-stage-to-orbit and return system with high development cost but very low recurring costs leading to very low life-cycle costs. The other approach proposes a very simplified expendable launch vehicle with extremely low both non-recurring and recurring costs, also leading to a very low life-cycle cost and early IOC (with low risk). This paper presents physical and cost-benefit arguments which result in recommending that the Ultra Low-Cost Expendable Launch Vehicle (LCELV) concept be seriously considered as the next logical step in achieving cost-effective access to space with minimum fiscal risk. Such an approach would, greatly reduce the cost of future U.S. commercial, civil and military space systems and would go a long ways towards restoring the international competitiveness of U.S. Launch Vehicles. 3. Design and Economic Considerations The basic issue is how can the U.S.A. best achieve it's national initiative to develop truly low cost, yet highly reliable space transportation, that is competitive/economically superior to the foreign off-shore launchers that have been taking away our launch vehicle business. Evidence of the severe decline in U.S. launch vehicle business is readily observed from data that show that the U.S. launch business has been reduced from 51% of the world's launches in the 1980s to only 34% in the 1990s. The U.S. currently does not offer low cost, highly reliable launch vehicles on a competitive basis compared to Ariane, Long March (China) and Proton (Russia). Only the MDA Delta is competitive but at the low end of the market (10,000 to 12,000 lb to LEO, maybe 4000 to 6000 lb to GTO). The U.S. has developed only three generations of launch vehicles during the entire history of the American Space Program. These are: 1. The IRBM, ICBM derived family of launchers from the period (Thor, Atlas, and Titan). 2. The Saturn 5, used for the Apollo and Sky Lab programs and then abandoned forever. 3. The Space Shuttle which has turned out to be a major financial burden for NASA and is currently barred from use (by a 1986 presidential decree) as a commercial satellite launcher. The U.S. has made several) attempts to begin new launch vehicle development programs with such initiatives as STAS, ALS, NLS and Spacelifter. But every one of these new launch system initiatives has failed because of the lack of an integrated national space transportation strategic plan. Both the Congress and the executive policy branches of government have rejected and withheld approval for every funding request from NASA for a new national launch vehicle development program for the last eight or nine years. Everybody associated with the U.S. Space Program agrees that a new modem launch vehicle is needed. Furthermore, every interested party agrees that these new launchers should be very low cost and competitive on a worldwide basis. The major problem that has prevented such a new launch vehicle program from moving forward is that none of the major organizations involved in establishing a new national launcher program can agree on an optimum conceptual approach for a new low cost vehicle.

4 The fundamental disagreement on the best approach for a low cost launch vehicle is whether the new vehicle should be reusable or expendable. Champions of the reusable concept assert that the lowest cost approach (they claim in fact, "the only way to achieve truly low cost") is to develop a highly reusable (up to 50 or 60 missions before overhaul), preferably single stage to orbit, vehicle that can be turned around for re-flight in a matter of hours with a minimum (three to ten people) maintenance/servicing crew. Proponents of this concept claim that the development through flight certification costs are acceptable and can be recovered at a reasonable amortization rate over the number of annual flights of the fleet of reusable launch vehicles (RLV). Proponents claim further that this approach will result in a launch capability of lb lo LEO for an average per launch cost * of $40 M. The NASA team that has been studying candidate new RLV concepts has identified the following potential reusable launcher configurations: Single stage to orbit: All rocket (SSTO-R) Two stage to orbit: All rocket (TSTO-R) Single stage to orbit: Air breathing plus rocket (SSTO-A/R) Two stage lo orbit: Air breathing plus rocket (TSTO-A/R) Of these general categories of RLV, only the first two (i.e., all rocket powered) were selected by NASA for more serious consideration. This was because the combined air breathing'' plus rocket power plant systems were considered to be far too complex and therefore more expensive lo develop at this time. Six high performance, high pressure engine concepts have been selected. These engine technologies are either of the tripropellant or full flow staged combustion (with advanced altitude compensation designs) type. None of these engine concepts has been developed past the early, partial feasibility stages and therefore all will present significant technology challenges that will most likely involve large financial investments lo achieve flight readiness. The six candidate rocket engine technologies currently under consideration for the RLV main propulsion system are summarized in Table 1. Table 1. Earth-to-Orbit Reusable Launch Vehicle (RLV) Engine Technology Candidates RD0120 Technology Derived Tripropellant Engine RS2000 Tripropellant Bell Annular Engine RD704 High Pressure, Oxidizer Rich Tripropellant, Single Thrust Chamber Engine AJ Full Flow L0 2 /LH 2 Staged Combustion Engine RS2100 Full Flow Staged Combustion L0 2 /LH 2 Bell Engine RS2200 Full Flow Staged Combustion L0 2 /LH 2 Aerospike Engine On the other hand, proponents of the low cost expendable launcher concept argue that the only way to achieve a truly low cost launcher is to have a two- or three-stage, all liquid rocket vehicle, designed with extreme low cost hardware and only moderate performance re requirements, so that the development through * Costs are in 1994 dollars.

5 qualification costs can be kept low enough that the resulting rate of amortization of development costs is acceptably low, insuring that this investment can be recovered after a reasonable number of flights. The assertions for RLV System propulsion needs, requirements and constraints, as recently established by the NASA/MSFC-led National Space Propulsion Team are reflected in their just completed reusable launch vehicle study, Reference 2, and are presented in Table 2. The initial assumptions adopted by this study team are shown in Table 3. It must be observed that many of these assertions for RLV needs, requirements, constraints and initial assumptions are strangely familiar; in fact, at closer look one might even conclude they are somewhat reminiscent of what was originally put forth for the Space Shuttle National Transportation System (NTS) concept (and upon which this concept was sold to Congress) back in the early 1970s. From the standpoint of overall launch vehicle capabilities, proponents of the RLV agree that a low cost reusable vehicle must meet the criteria delineated in Table 4. The proponent of the low cost expendable launch vehicle argue that high performance and corresponding low GLOW are not required for low cost and that performance goals that are too high will, in fact, lead to high cost launchers even if they are expendable. Furthermore, for expendable launchers, the hardware is 80% of the total launch cost with only 20% required for operations (see Figure 4 for a breakdown). Of the 80% for hardware, approximately 40 to 50% of the cost is for propulsion and 30% is for the structure. Therefore, the key to low-cost expendable launchers is to design and develop low-cost propulsion systems (especially the rocket engines) and low-cost structures and tanks which is illustrated in Figure 5. There is a direct correlation on propulsion costs as a function of parts and interface counts and chamber pressure as shown in the cost versus chamber pressure chart. The main cost driver for propulsion system costs is the operating chamber pressure, Pe. For chamber pressures below 1,000 psia, there is a wide range of options ranging from emphasis on minimum weight designs such as the F-1, J-2, and R- 10 to the LCE which is designed for minimum cost with far fewer parts and interfaces compared to the previously mentioned designs. Thus there is a wide range in development and recurring costs resulting from lower to higher man-hour efforts due to a wide variation in complexity. As one begins to strive for maximum performance by operating at chamber pressures well above 1,000 psia, for example, the STME at Pc = 2,250 psia and the SSME at Pe = psia, the acceptable designs are restricted to a narrow band of highly complex and hence costly designs. Within a fixed chamber pressure design class, the recurring cost of an engine scales approximately as the square-root of the thrust level since the thrust is proportional to the throat area. In the case of varying the chamber pressure there is an exponential increase in the recurring cost of the enging at a fixed thrust level as the chamber pressure is increased beyond about 1,000 psia. Figure 5A is a plot of the heuristically-derived formula for the first unit cost, C (in millions of dollars) as a function of thrust level, F (in pounds) and chamber pressure, Pe (in psia) in 1995 dollars. The formula is expressed as: 0.5 ( Pc /850) C = 2,000( F) e All engine thrust levels have been normalized to F = 500,000 pounds. This meant scaling F- 1 down from 1,500,000 pounds, J-2 up from 250,000 pounds, French Vulcain up from pounds, and RL-10 up from 30,000 pounds of thrust. The dashed curve represents the spread in cost of the various designs at the low pressure end of the spectrum to a near vanishing of the spread at the SSME level of Pc = 3,260 psia. Figure 5B is a plot of the LCE designs fixed at Pe = 300 psia for a thrust range from F = 40,000 pounds to F = 1,000,000 pounds. The design points ate TRW design cost estimates. Using a 90% learning curve would yield an average unit cost of 0_5C for one hundred manufactured engines. Figure 5C shows the variation in first-unit costs for LCE ranging in design thrust levels of 40,000 to 1.000,000 pounds with chamber pressures increasing from 300 psig to 1,500 psig. Finally, the arguments for low-cost expendables emphasize that low-cost propulsion systems and low-cost structures must meet the criteria listed in Table Conclusions and Recommendations As a result of the performance considerations, single-stage to low-earth orbit vehicles utilizing LOX- Hydrogen must achieve mass fractions of less than γ = to obtain payload ratios of P = 50 (2% payload). In recent design studies by NASA/Langley Research Center, two typical vehicles capable of placing 25,000 payloads in low-earth orbit and return require gross lift-off weights of 3,115,926 lbs at a payload ratio of P =

6 125 (0.85% payload) or in the case of a LOX-RP-Hydrogen (triprop) design a gross lift-off weight of 2,454,397 lbs or a payload ratio of P = 98 (1.02% payload). These designs require mass fractions of y = and y = 0.086, respectively, in the actual designs which is close to the theoretical results of Section 2. Both of these designs have winged entry and landing gears and require, typically. seven high-pressure engines. The two-stage simplified expendable designs, with either LOX-Hydrogen or LOX-RP, require one to four lowpressure booster engines (depending upon the configuration) of far less complexity and cost than the highpressure engines. As illustrated in Figure 6, sensitivity to changes in the mass fractions is very great for the single-stage case with respect to changes in the payload ratio with a change in P of 20 points plus or minus for a change in mass fraction of plus or minus Similarly, as shown in Figure 6, a change in the average specific impulse of plus or minus 1.0 seconds results in an opposite change in the payload ratio, P, of minus or plus 9. Thus an increase of in the mass fraction can be balanced by an increase of the specific impulse of 2.2 seconds or vice versa. The sensitivity to changes in mass fraction and/or specific impulse, at a payload ratio of P = 100 (1 % payload) as reflected in the Langley designs, reflect the closeness to a cut off in payload capability for any deficiencies in the design assumptions with respect to propulsion performance and component weights. The multistage versions, allowing for greater performance variations while at the sane time allowing for the use of much simpler designs, reflect a potential for low development and recurring costs. For the two-stage LOX-Hydrogen design point, payload percentages are nearly 80% of what can be achieved by the ultimate case of "infinite" staging as discussed in Section 2 and shown in Figure 6. The single-stage to low-earth orbit, reusable, as represented by relatively detailed design efforts by NASA/Langley, results in payload ratios of approximately P = 100. This means that less performance causes the payload ratio to begin increasing exponentially to an infinite payload ratio or zero payload. This can be seen in Appendix A where the denominator for the single-stage equation for the payload ratio. P. is the difference of two terms, one involving the specific impulse, 1. and the other term is the mass fraction, y. An increase in y with a decrease in 1 causes any positive value of the denominator to approach zero which makes P approach infinity. Since the single-stage case is near this point. considerable efforts (cost) need to be expended to hold onto to a design with a reasonable payload ratio which can nonetheless be a robust enough design to be reused many times with very low recurring costs. Current efforts are concentrating on Option b) (SSTO-R, TSTO-R) by NASA and Option d) (EELV) by AFSMC. Option a) (LCELV), which by the arguments in the previous discussions has the best chance to make available near-term launch-cost reductions, has been ignored or dismissed by the relevant government agencies and space industries and needs some serious reconsideration. Current estimates for DDT&E costs for Option b) have a consensus industry/ government estimate of the order of $10-$20 billion (Reference 3), whereas Option a) has been estimated to be about $0.5-$1 billion (Ref. 4). Amortizing these amounts over one hundred launches adds $100-$200 million to the recurring cost for each Option b) launch versus $5-$10 million to the recurring cost for each Option a) launch. Recurring costs (including amortization of DDT&E, production and launch support) for the 25 Klb payload-to-leo for the Option a) vehicle would range from $20 to $25 million per launch. Proponents for Option b) have estimated forty launches per year for up to thirty years which would reduce the amortization of DDT&E by nearly an order of magnitude (Le., to approximately $10-$20 million per launch). However, refurbishing and/or replacement costs could far exceed expectation and negate the assumed cost/benefits for rapid amortization of DDT&E because of very low recurring costs. In other words, far fewer flights may be achieved by each SSTO-R with concurrently high refurbishing costs to keep it flying reliably. A more recent cost evaluation (Reference 5) by NASA estimates a procurement of five SSTO-Rs at a cost of $2B each with a target of 100 flights per vehicle or $800 per pound of payload to LEO. At a flight rate of 40 per year, the cost for Operations and Support adds another $800 per pound to LEO or a total of $1,600 per pound to LEO. A rationale for reducing these costs by an order of magnitude with such things as flight reuse greater than 1,000 are presented. However, by using he original estimate of 100 flights per vehicle and amortizing a nominal development cost of $15B over 500 flights for the fleet of five vehicles adds $30M per flight or gives a total of $2,800 per pound to LEO. The costs are summarized in Table 6. A similar summary of costs for a typical two-stage LCELV (using LOX-Hydrogen) is shown in Table 6. The development cost is estimated at $600M with an average production cost of $9.5M assuming a 90%

7 learning curve for a procurement of 100 units. Operations and Support cost per flight are $9.5M and adding $6M per flight for amortization of the development cost, yields a cost per launch of $25M or a total of $ 1,000 per pound to LEO. The numbers are also summarized in Table 7 with a more detailed breakdown. With low development costs and low recurring costs (estimated to be one-fourth to one-fifth those of the current expendable systems), the LCELV has the potential for the lowest lifecycle costs because relatively low DDT&E can be economically amortized over one hundred launches while at the same time recurring costs will be low. Furthermore, the LCELV can be developed for near-term use with IOC as early as three years from program start. While payload requirements, to LEO. have been set to between 20K - 25K lbs, any future desires for much larger payloads (i.e., 100K lbs) would be beyond any reasonable gross weight for the SSTO- R since the nominal payload ratio of about 100 would yield gross lift-off weights in excess of 1OM lbs. In the case of LCELV, with a payload ratio of about 25, the gross lift-off weight would be about 2M lbs which is no more than the current SSTO-R concepts. Table 4.- Criteria Low Cost Reusable Launch Vehicle (RLV)Propulsion System Minimum Number of Components High Component Reliability at Specified Confidence Level (+) Mean Time Between Overhauls (+) Number of Toxic Fluids (-) High Design Life Margin Ease of Maintenance Number of Different Propulsion Systems (-) Degree of Self Test (+) Turnaround Cost/Schedule/Complexity (-) Minimum Number of Interfaces (System, Component, Ground) (System to Ground, System) Number of Inspection Points (-) Number of Check-outs Required (-) Number of Systems/Components Requiring Monitoring (# Hazards) (-) Minimum Components That Have to be Overhauled Number of Criticality 1 Failure Modes (-) Controllability with Engine Out High Pressurization Components (-) Ease of Loading and Conditioning Number of Different Fluids in System (-) Number Purges Required (-) Number Bleeds (-) High Average Specific Impulse (+) High Installed Propulsion System Specific Thrust (+) Years to Payback (-) Hardware Recurring Cost (-) Quantity of Pollutants (-)

8 References Table 5. Criteria for Low Cost Expendable Launch Vehicles Use simple mocoque or low-cost composite structures instead of complicated honeycomb, isogrid, etc. Use low engine operating pressures ( psia) Use liquid propulsion two stages for earth-to-orbit launcher Engines should be passively cooled (using low cost, high durability ablative liners) Use pressure-fed system for small launchers and simple, foil bearing, hydrodynamically lubricated and cooled, pumps for larger launchers Minimize engine and system parts count, i.e., reduce parts count on the order of two orders of magnitude from that of the Space Shuttle Reusable launch vehicles will, of major necessity, require very high performance propulsion systems and lightweight structures to be at all practical which will lead to very high development costs that can never be amortized at any kind of reasonable rate. Use simple propulsion system to minimize development through qualification costs Design rocket engines that avoid combustion instabilities during development and qualification Use a simple gas generator engine cycle to drive pumps Avoid complex staged combustion or expander cycle engines that operate at very high pressures (i.e., >2500 psia) because these conditions directly drive costs Avoid winged fly-back boosters Avoid rocket powered vertical launchers like DC-X that land under rocket power when returning from orbit Reusable launchers will have very high cost thermal protection systems (TPS) which will also greatly drive up cost of vehicle development Use reasonable low-cost materials to the maximum extent possible for all systems/subsystems 1. "National Space Transportation and Support Study ," by the Joint DoD/NASA Transportation Technology Team. Preliminary Report May "Reusable Launch Vehicle (RLV) Propulsion Technology Implementation Plan," NASA Space Propulsion Synergy Team, October 30, "Access to Space Study" (Summary Report), Office of Space Systems Development, NASA Headquarters. January, "Low-Cost Launch Vehicle Study," TRW Systems Group for NASA Headquarters. Final Report (Contract No. NASW-1792). June 23, "Highly Reusable Space Transportation." John C. Mankins, NASA Advanced Concepts Office, Briefing February 15, 1995.

9 AEROSPACE AMERICA/JANUARY 2006 Crew Exploration Vehicle Configuration Following the completion of the ESAS, NASA chose a blunt-body capsule shape as the CEV reentry vehicle. T o make maximum use of the Apollo research and development work, the same slant-angle shape of the Apollo command module was chosen. The CEV command module, however, would be a larger cone, 5.5 m wide at the base, as compared to the Apollo CM 3.9-m width. The capsule would be partially reusable, with an expendable heat shield and an expendable launch escape rocket on top for launch aborts. The CEV CM would come in different versions, from all-cargo configurations to versions that would carry three to six crewmembers, depending on the mission. Beneath the command module the CEV would have a service module, with a LOX/methane engine for propulsion and solar arrays for electrical power. The command module, service module, and launch escape system would be mounted on top of a four-segment solid fuel booster rocket derived from the space shuttle booster, but with a new liquid upper stage powered by a modified space shuttle main engine. The booster would be recoverable; the upper stage and service module would be expendable. Following completion of either a low Earth orbit mission or a lunar flight, the CEV command ship would reenter the atmosphere and land on a hard surface, lowered by parachutes and some possible combination of retrorockets or airbags. The capsule would also be designed for recovery at sea in the event of a launch or mission abort. Details of the configurations will be discussed in Part II.

10 AEROSPACE AMERICA/JANUARY 2006 Various concepts of launch vehicle configurations were put forward, for the different roles the CEV would play. The early versions of the CEV will focus on delivering humans to the Moon, first for brief sorties, then for extended stays. The CEV, its launchers, its propulsion systems, and the accompanying landers will take the U.S. manned space program in a whole new direction

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12 Concept image of the Ares V earth departure stage in orbit, shown with the Crew Exploration Vehicle docked with the Lunar Surface Access Module. (NASA/MSFC)

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18 HOPPER: a possible future space transport system In its HOPPER concept, EADS SPACE Transportation has gone for an autonomous transport system that is noted for its high degree of reusability and considerably lower mission costs. The unmanned HOPPER will be launched horizontally on a skid sled running on a four kilometre long track. With a length of more than 50 metres and a wing span of more than 27 metres, the vehicle itself is quite compact. Due to its features, Hopper is intended to transport payloads to orbit at considerably lower cost than conventional transport systems. HOPPER is designed to carry payloads of up to 7.5 tons - as a rule a satellite on an upper stage booster. At an altitude of 130 kilometres, HOPPER will deploy the payload from its tail and then automatically return to Earth while the payload proceeds independently to its destination in geostationary or low-earth orbit, such as the ISS for instance. HOPPER will be launched from the European Space Centre in Kourou, French Guiana, and, after a flight across the Atlantic Ocean, will land on European territory (e.g. islands in the Atlantic). In comparison to the Shuttle, which orbits the Earth several times before re-entry, HOPPER is indeed a more efficient system. HOPPER can be transported back by ship. Should the ESA decide in favour of the HOPPER concept the space vehicle will be ready for use by the year REUSABLE SUB-ORBITAL HOPPER DESIGN