The fundamental ideas underlying reusable vehicle concepts, such as the reusable booster system (RBS), are not new, and many significant activities have been conducted in both the United States and overseas that are aimed at realizing the potential gains that come with reusability. In this appendix, three large U.S. efforts are briefly reviewed and lessons learned are discussed.
The space shuttle is the only partially reusable launch system that has successfully been used over an extended time period (USSR’s Buran, which was unmanned, was only flown to orbit once). The orbiter is the principal reusable stage of the space shuttle; its solid rocket boosters were recovered and refurbished and its external tank was expended. The orbiter is a manned vehicle that goes into orbit, stays there for several weeks, then reenters and glides back for landing on a runway.
To comply with the 1969 Space Task Group recommendation that NASA focus on a reusable launch system, it was determined that the keys to space shuttle acceptance were to significantly reduce cost and obtain broad support from all involved government agencies and organizations. As a result of the phase A studies, a fully reusable two-stage concept was the best choice for lowest operational cost per flight. However, it was apparent that concurrent development of both an orbiter and a booster vehicle would be expensive, since peak-year development funding had become a concern. In 1971, the Office of Management and Budget direction constrained the space shuttle overall development cost and limited the annual funding for development to approximately half that estimated to be required for the fully reusable two-stage configuration. These budgetary constraints forced NASA to consider other alternatives, which ultimately resulted in the partially reusable configuration known today.
There were major drivers that impacted NASA’s ability to achieve the objective of significantly reducing space shuttle’s operational cost:
• Political decisions heavily influenced the development process;
• Budgetary constraints drove the design;
• Abort capability with engine out (two engines provided best performance and lowest cost);
• Department of Defense (DOD) cross-range requirements drove the configuration and added complexity;
• Continuous additional requirements were introduced during the development process; and
• The complexity of the configuration and its resulting sensitivity and required interactivity provided design challenges, which affected the operational complexities and flight constraints.
Even if the original, totally reusable two-stage human-rated space shuttle system had been properly funded from the outset, it is doubtful that the initial low cost goals would have been met. This conclusion is further supported by what is now recognized as an unrealistically projected high launch rate.
The orbiter’s design was performance driven. It needed to achieve low Earth orbit, provide life support for its crew for several weeks, reenter Earth’s atmosphere, and glide to a safe landing. The ability to reserve some orbiter weight during its development to make its ground turnaround operations more efficient was not possible. As a result, the anticipated flight rate was never achieved because ground turnaround operations were extensive and very costly.
Maintenance issues associated with the orbiter’s reentry thermal protection system (TPS) tiles made rapid ground turnaround impossible. After every flight, each of the orbiter’s more than 27,000 tiles had to be individually inspected for damage and adhesion and manually replaced, if necessary.
The orbiter’s three main reusable engines (space shuttle main engines, or SSMEs) also required inspection after landing. Because of access and interference issues with other orbiter components, SSME’s were removed and replaced after every space shuttle orbiter mission, beginning with operational mission STS-6. Off-orbiter engine processing during the operational phase was done for several reasons. Checkout of the SSMEs required power-up of electrical and fluid subsystems, which interfered with other orbiter processing. Also, access was often needed in the orbiter’s engine compartment to other main propulsion system components, and the installed SSMEs restricted access and often resulted in serial processing work in the aft compartment. It was easier (and more efficient) to remove and replace engines than to accommodate these other needs with engines in place. Orbiter and engine processing work could be accomplished in parallel with removal of the engines, thus any unforeseen contingency processing would not impact other components.
This orbiter ground processing experience resulted in the development by NASA of operability/maintainability design considerations for next generation reusable launch vehicles.
NATIONAL AEROSPACE PLANE HISTORY
The National Aero-Space Plane (NASP) was a project jointly funded by NASA and DOD to create a single-stage-to-orbit spacecraft.1 The NASP concept evolved from the “Copper Canyon,” the Defense Advanced Research Projects Agency (DARPA) project, running from 1982 to 1985. The vehicle was planned for a crew of two and was meant to serve access-to-space missions. In 1990, McDonnell Douglas, General Dynamics, and Rockwell International formed a national team to develop a demonstrator vehicle, the X-30, to deal with the technical and budgetary issues.
There were six technologies considered critical to the success of the project; three of them were related to the propulsion system. The X-30 was intended to use an engine that could shift from a low-speed propulsion system to a scramjet system as the vehicle ascended; it would burn liquid hydrogen fuel with oxygen taken from the atmosphere. An auxiliary LO2/LH2 rocket engine was used to augment the scramjet engine at very high speeds and for propulsion needs in space.
1 B.W. Augenstein and E.D. Harris, The National Aerospace Plane (NASP): Development Issues for the Follow-On Vehicle, R-3878/1-AF, RAND Corporation, Santa Monica, Calif., 1993; Encyclopedia Astronautica, X-30, available at http://www.astronautix.com/lvs/x30.htm, accessed on October 9, 2012.
The scramjet engine was a key to the NASP multi-cycle engine. In a scramjet the onrushing hypersonic air is compressed and passed into a combustion chamber in which hydrogen is injected and burned by the hot, compressed air. The exhaust is expelled through a nozzle creating thrust. The efficient functioning of the engine depends on the aerodynamics of the airframe, the underside of which serves as the air inlet and the exhaust nozzle. Design integration of the airframe and engine are thus critical to the success of the design.
Other enabling technologies include the development of materials that would maintain structural integrity at very high temperatures, sometimes in excess of 1,800°F. The enormous heat loads associated with hypersonic flight have required the development of active cooling systems and advanced heat-resistant materials.
Using atmospheric oxygen instead of tanked LO2 for the majority of the mission, the NASP airbreathing engines were projected to have a specific impulse that is approximately 3 to 4 times that of a LO2/LH2 rocket engine. This enables a single-stage-to-orbit (SSTO) vehicle with propellant mass fraction of approximately 76 percent to achieve orbital speeds. So the principal technical challenge was to achieve the high performance of the airbreathing engines, while limiting the impact of the inert mass of these engines and the additional mass of the thermal protection system required to protect the vehicle during its ascent to orbit.
Despite significant progress in structural and propulsion technology, the program had substantial hurdles to overcome. DOD wanted it to carry a crew of two, plus a small payload. The demands of being a man-rated vehicle operating as a SSTO vehicle made the X-30 more expensive, larger, and heavier than is required for a demonstrator vehicle.
As a SSTO vehicle with a turnaround time of 24 hours or less, proponents initially viewed the X-30 as leading the way to faster, cheaper access to low Earth orbit. What became obvious was that the claims made for NASP as a space launch vehicle were similar to the initial claims made for the space shuttle in the early 1970s. The assertions that NASP would have airplane-like operating characteristics were assumptions, not conclusions based on detailed analysis.
NASP never achieved flight status and finally fell to budget cuts in 1993. But it also was cancelled because of its severe technological overstretch. Although the X-30 program never came near to building significant hardware, NASP development work contributed in an important way to the advancement of propulsion technology and high-temperature materials as well as materials capable of tolerating repeated exposure to extremely low temperatures (the cryogenic fuel tanks). By 1990, NASP researchers had realized significant progress in titanium aluminides, titanium aluminide metal matrix composites, and coated carbon-carbon composites. Moreover, government and contractor laboratories had fabricated and tested large titanium aluminide panels under approximate vehicle operating conditions, and NASP contractors had fabricated and tested titanium aluminide composite pieces.
When NASP was cancelled, the government admitted to making a $1.7 billion investment in the National Aerospace Plane, but parts of the research and development were classified, and the official costs may have been higher.2
Lessons learned from the NASP program include the following:
• SSTO vehicle technologies using airbreathing engines are beyond the existing state of the art;
• Aerothermodynamics of sustained flight at high hypersonic speed within the atmosphere creates significant challenges with regard to materials and thermal management; and
• Propulsion technology should be independently matured prior to initiation of large-scale vehicle development programs.
The X-33 VentureStar was intended to be a one-third-scale prototype of a Reusable Launch Vehicle (RLV), designed to significantly lower the costs of launching payloads. VentureStar was a very ambitious effort by NASA and Lockheed Martin to develop a fully reusable, vertical launch and horizontal landing, manned, (SSTO vehicle. To achieve SSTO capability, VentureStar needed to incorporate high thrust and high-specific-impulse propulsion
2 R. Launius, After Columbia: The Space Shuttle Program and the crisis in space access, Astropolitics 2:277-322, 2004.
and have a mass fraction higher than 0.88. During development of a SSTO, any small increase in vehicle weight directly offsets payload capability, requiring either a very strict weight reduction activity (costly) or vehicle up-scaling (even more costly). If the SSTO vehicle’s mass fraction falls below 0.88, its weight for delivery of a given payload to orbit grows asymptotically as the vehicle becomes unable to lift even itself into orbit. Today’s most efficient expendable upper stage (Centaur) has a mass fraction of 0.90. Making a stage larger helps, but adding thermal protection for reentry, aerodynamic surfaces for lift and control, landing gear, and life support for manned use makes meeting the 0.88 minimum mass fraction requirement extremely difficult. X-33 was to address and demonstrate the many technology issues associated with meeting this requirement.
Significant advances were made with the thermal protection system developed by B.F. Goodrich, which also served as the aerodynamic surface of the vehicle. The Rocketdyne XRS-2200 Linear Aerospike main engines were on target to become the next generation of liquid fueled propulsion systems. Both of the conformally shaped LH2 and LO2 propellant tanks were initially to be composite, but after engineering objections, the smaller LO2 tank was successfully manufactured with aluminum-lithium (Al-Li). Management insisted, however, that for X-33 to be a useful technology enabler, the multi-lobed LH2 tank had to be made of composites.
Development problems were encountered with the LH2 tank, along with Rocketdyne’s decision to use Narloy-Z (a heavy copper alloy), which drove changes to the flight control surfaces. These major issues (and other smaller ones) resulted in substantial cost increases and schedule delays. The composite hydrogen tank failed during testing, which resulted in program cancellation in 2001. X-33’s development cost a reported $1.5 billion over 5 years; the vehicle was approximately 40 percent complete at cancellation, and its test flight facilities at Edwards Air Force Base were ready. If the LH2 tanks had been manufactured of Al-Li instead of a composite as proposed by the X-33 engineering team, they would reportedly have been lighter than the composite tank and, therefore, successful. Unfortunately, all the successful new technology was laid to rest along with the death of the X-33, and the opportunity to gather useful flight information that would have been applicable to a reusable booster was lost.3
In addition to the three large programs discussed above, there have been and continue to be numerous smaller efforts in the United States and internationally aimed at development of reusable launch vehicles. Within the United States, these include the X-20 Dyna-Soar, X-34, Delta Clipper, Kistler K-1, DreamChaser, Prometheus, and Blue Origin New Sheppard. Internationally, these include the HOTOl, Skylon, Sanger, and Buran. A detailed investigation of these programs was beyond the statement of task for this committee.
A number of important differences exist between the three reusable launch vehicle programs described above as compared to the RBS concept, including the following.
• The three RLV programs described above were intended for manned operations, which imposes numerous subsystem requirements that add to the inert launch vehicle mass. The RBS concept is not intended to be human rated, so it will not be burdened by this additional inert mass requirement.
• The previous RLV programs involved reusable vehicles that were launched into space and needed to survive orbital reentry conditions. In the base of the RBS concept, the maximum Mach number of the reusable booster is between 3 and 7, so the requirements for a thermal protection system are greatly diminished.
• The space shuttle used a reaction-control system that operated with toxic propellants, which greatly complicated operations and significantly impacted turn-around time. As a new design, the RBS concept development can pursue use of non-toxic propellants for the attitude-control system, which would provide significant operability advantages.
• Most RLV programs were based on the use of hydrogen as the fuel. (An exception is the Kistler K1 project, which was based on the use of LO2/RP propellants.) The RBS concept is based on the use of a hydrocarbon fuel for the reusable booster. This higher density fuel will likely result in a more structurally and aerodynamically efficient vehicle configuration.
3 R. Launius, After Columbia: The Space Shuttle Program and the crisis in space access, Astropolitics 2:277-322, 2004.
Also, if the RBS development were to be pursued using the staged development approach as recommended by this committee, each of the technology risks would be matured to a much higher technology readiness level prior to committing to full-scale vehicle development and flight certification, as compared to the above RLV programs. This approach contrasts dramatically with the manner in which technology risks were addressed in the previous RLV programs. In these programs, the major technology risks (i.e., the reusable SSME and solid rocket motor for the space shuttle; the combined-cycle engine for NASP, and the ultra-lightweight structure with hydrogen fuel for the X-33) were all developed in parallel with the full-scale vehicle development.
With these considerations, there are clear differences between the RBS concept and previous RLV programs. And, while efficient reusability remains an elusive goal for launch, the committee believes that the RBS concept represents a logical compromise between fully reusable and fully expendable systems that is technically achievable in a well-structured program. So, while the committee does not believe that the business case of the RBS concept can be closed at this time, it is important to continue to mature the underlying technologies.