4
Rocket Propulsion Systems for Access to Space

INTRODUCTION

The U.S. Space Transportation Policy calls on the Secretary of Defense (SECDEF), in coordination with the National Aeronautics and Space Administration (NASA), to be responsible for assuring access to space for critical national security, homeland security, and civil missions. Assured access to space is defined as “a sufficiently robust, responsive, and resilient capability to allow continued space operations, consistent with risk management and affordability” (NSPD, 2005, p. 4). Such access will require maintaining a viable industrial and technology base.

The SECDEF is also called upon, before 2010, to begin a fundamental transformation in the U.S. capability for “operationally responsive access” to and use of space “that dramatically improves the reliability, responsiveness, and cost of access to, transport through, and return from space.” This requires a sustained technology development program to pursue research and technology development in in-space transportation capabilities, including automated rendezvous and docking and the ability to deploy, service, and retrieve payloads or spacecraft in Earth orbit.

The U.S. Space Transportation Policy calls for development of requirements, concept of operations, technology roadmaps, and investment strategy for next-generation space transportation capabilities within 2 years (NSPD, 2005).

The Air Force Space Command’s (AFSPC’s) Strategic Master Plan FY06 and Beyond states as follows: “AFSPC will sustain and modernize its current satellite and launch operations into the far term, when it will transition to advanced capabilities” (AFSPC, 2003, p. 29).1

The Air Force’s overarching need to have responsive access to space and to operate effectively in space under all realistic scenarios will demand the establishment of requirements for (1) strategic and responsive spacelift total systems, for (2) responsive on-board propulsion systems in space and for (3) return from space. Transformation in access to space or in-space operations will be achieved only as a result of using a total systems engineering process incorporating mission success over the committed life of the system as the primary criterion when selecting among options for a required system’s architecture and elements. The evolution of such a system engineering program, and the validation of trade-off parameters using the supercomputing capabilities available today, would provide a powerful and objective quantitative tool to define and evaluate low-risk, cost-effective total system concepts for strategic and operationally responsive spacelift (ORS) and for in-space operations. “Mission success” is the most effective selection criterion in a total systems engineering process to establish an overall architecture and all the elements of a system. It can be defined as achieving the functional result we want, when we want it, for the price we committed to, and within the risk level profile we accepted for the

1

For responsive spacelift, the Air Force Space Command’s Strategic Master Plan FY06 and Beyond defines transformational capabilities as focused on rapid response, affordability, and payload capacity for warfighter operations (AFSPC, 2003).



The National Academies | 500 Fifth St. N.W. | Washington, D.C. 20001
Copyright © National Academy of Sciences. All rights reserved.
Terms of Use and Privacy Statement



Below are the first 10 and last 10 pages of uncorrected machine-read text (when available) of this chapter, followed by the top 30 algorithmically extracted key phrases from the chapter as a whole.
Intended to provide our own search engines and external engines with highly rich, chapter-representative searchable text on the opening pages of each chapter. Because it is UNCORRECTED material, please consider the following text as a useful but insufficient proxy for the authoritative book pages.

Do not use for reproduction, copying, pasting, or reading; exclusively for search engines.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs 4 Rocket Propulsion Systems for Access to Space INTRODUCTION The U.S. Space Transportation Policy calls on the Secretary of Defense (SECDEF), in coordination with the National Aeronautics and Space Administration (NASA), to be responsible for assuring access to space for critical national security, homeland security, and civil missions. Assured access to space is defined as “a sufficiently robust, responsive, and resilient capability to allow continued space operations, consistent with risk management and affordability” (NSPD, 2005, p. 4). Such access will require maintaining a viable industrial and technology base. The SECDEF is also called upon, before 2010, to begin a fundamental transformation in the U.S. capability for “operationally responsive access” to and use of space “that dramatically improves the reliability, responsiveness, and cost of access to, transport through, and return from space.” This requires a sustained technology development program to pursue research and technology development in in-space transportation capabilities, including automated rendezvous and docking and the ability to deploy, service, and retrieve payloads or spacecraft in Earth orbit. The U.S. Space Transportation Policy calls for development of requirements, concept of operations, technology roadmaps, and investment strategy for next-generation space transportation capabilities within 2 years (NSPD, 2005). The Air Force Space Command’s (AFSPC’s) Strategic Master Plan FY06 and Beyond states as follows: “AFSPC will sustain and modernize its current satellite and launch operations into the far term, when it will transition to advanced capabilities” (AFSPC, 2003, p. 29).1 The Air Force’s overarching need to have responsive access to space and to operate effectively in space under all realistic scenarios will demand the establishment of requirements for (1) strategic and responsive spacelift total systems, for (2) responsive on-board propulsion systems in space and for (3) return from space. Transformation in access to space or in-space operations will be achieved only as a result of using a total systems engineering process incorporating mission success over the committed life of the system as the primary criterion when selecting among options for a required system’s architecture and elements. The evolution of such a system engineering program, and the validation of trade-off parameters using the supercomputing capabilities available today, would provide a powerful and objective quantitative tool to define and evaluate low-risk, cost-effective total system concepts for strategic and operationally responsive spacelift (ORS) and for in-space operations. “Mission success” is the most effective selection criterion in a total systems engineering process to establish an overall architecture and all the elements of a system. It can be defined as achieving the functional result we want, when we want it, for the price we committed to, and within the risk level profile we accepted for the 1 For responsive spacelift, the Air Force Space Command’s Strategic Master Plan FY06 and Beyond defines transformational capabilities as focused on rapid response, affordability, and payload capacity for warfighter operations (AFSPC, 2003).

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs program. Improvement in mission success achieves that functional result for a lower price and with less risk. For example, when carrying out systems engineering for access-to-space missions, considerations that must be specified for a “total system” that will accomplish the mission are launch vehicle configuration (number of stages, reuse), launch locations (fixed or mobile, including from high-altitude aircraft), facility and logistic requirements, operations concepts (payload integration on launch stand or pre-integrated, attachable payload modules), technology validations still required, full development schedules, life-cycle costs, and industrial support viability. Only the unbiased application of such a systems engineering program could provide a credible basis for specifying requirements like number of stages, propellant, and reusability or flexibility of launch locations. The systems engineering program would incorporate the ability to quantify relative risk and allow the selection of system options. Propulsion system requirements and basic configurations for propulsion subsystems would be an output of the process. Then, the design criteria can be set that, when satisfied, will assure that a subsystem performs as it should. Identifying missing or unvalidated design criteria associated with the selected propulsion systems for ORS would define the critical gaps in the technology base. AFSPC’s evolving space operations plan encompasses both low- and high-tempo operations. Low-tempo operations generally involve the planned placement and support of strategic and capital assets. Those assets are usually large and are placed in various orbits, often in geostationary orbit (GEO). High-tempo operations involve a quick response to a perceived threat buildup and may involve intensive launch and in-space operations lasting from days to months. DoD’s need for quick, responsive space launch under numerous scenarios drives its requirements for responsive spacelift, responsive stages in space, and responsive platforms with onboard propulsion systems. Those requirements, in turn, flow down via extensive systems engineering into a broad and demanding set of new requirements for propulsion systems for access to space and for in-space operations. Figure 4-1 presents a roadmap setting out the Air Force’s plan to sustain and modernize its current satellite and launch operations into the far term, when it envisions transformational access-to-space capabilities. The roadmap shows the near-term phaseout of Atlas II/III, Delta II, and Titan IV from the Air Force launch vehicle operational mix. Therefore, the AFSPC’s Strategic Master Plan FY06 and Beyond (hereinafter referred to for convenience as SMP FY06) to sustain and modernize current satellite and launch operations into the far term will be implemented primarily using the Atlas V and Delta IV vehicles along with several smaller and medium-lift vehicles that are used by the Air Force but not shown in Figure 4-1. The planned introduction and evolution of new small and mid-size launch vehicle capabilities are mapped in the Responsive Spacelift region of the roadmap. The new small vehicles planned for demonstration under the Force Application and Launch from the Continental United States (FALCON) program of the Defense Advanced Research Projects Agency (DARPA) are aimed at short-response launch times and low-marginal-cost launches. The Air Force intends to achieve the DARPA goals by using innovative but conventional rocket propulsion system elements, simple configurations, high safety margins against critical failure modes, and rapid-installation, standardized modules containing pre-checked-out payloads. As shown in Figure 4-1, there is no current plan to replace either the Atlas V or the Delta IV until some time well beyond 2020. There are good reasons why this is realistic. The capabilities required by the Air Force to deliver a mix of large payloads into the near-Earth region of space under the low-tempo operations scenario will be satisfied into the far term (beyond 2020) by modest continued evolution of the Atlas V and Delta IV configurations and upgrades of elements of their propulsion systems. The committee did not find any technologies currently in development or expected to be validated during the planning period for liquid- or solid-propellant, all-rocket (non-combined-cycle) propulsion systems for space access that would demonstrate enough improvement in performance or reduction in operational risks and costs necessary to justify the huge cost of a new centerline launch vehicle of the evolved expendable launch vehicle (EELV) class.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 4-1 DoD space transportation roadmap. SOURCE: Knauf (2005). Continued evolution of the technology used in materials, pumps, injectors, and thrust chambers and to improve engine thrust/weight, margins to failure modes, and other parameters of propulsion system elements will not enable truly transformational vehicle alternatives for any launch vehicles, large or small, in the near to medium term. Transformational technologies that can be envisioned for the far term are combined-cycle air/rocket engines that minimize the amounts of tanked oxidizers and/or very energetic, but stable and cost-effective new rocket fuels delivering low-molecular-weight combustion products. Such technologies, which might improve overall mission success by 25 to 50 percent, might justify investment in truly next-generation, large access-to-space vehicles to replace the Atlas V and Delta IV vehicles. Finding 4-1. The committee does not believe that the Air Force will be able to reliably and cost effectively transform U.S. military space transportation capabilities by focusing on pushing high-thrust rocket propulsion technologies to their limits. Even if the total systems optimization process is objectively carried out, the technologies it selects are unlikely to be (and need not be) transformational in themselves. It is more likely that any transformational access to space achieved during the planning period will be the result of creative total system architectures. Focusing Air Force resources on identifying the gaps in the critical design criteria for total systems-defined rocket propulsion elements will be crucial to success of the AFSPC Strategic Master Plan FY06 and Beyond.2 Recommendation 4-1. The Air Force should place a high priority on developing an integrated total-system engineering process using quantitative life-cycle mission success as the selection criterion for near-term, highly leveraged engineering technology funded by the Air Force. This process is crucial to 2 There are a number of new propulsion technologies that do in fact have the potential of directly enabling transformation of in-space rocket propulsion systems performance. They are discussed in Chapter 5, Propulsion Systems for In-Space Operations and Missiles.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs defining justifiable total system architectures, rocket propulsion systems requirements, and critical technologies for military space transportation to support the Air Force Space Command’s Strategic Master Plan FY06 and Beyond. CURRENT CAPABILITIES OF LARGE LAUNCH VEHICLES Delta IV Family of Vehicles As shown in Figure 4-2, the Delta IV family of two-stage launch vehicles utilizes a common 5-m-diameter first stage powered by a single rocket engine (RS-68) operating on liquid oxygen (LOx) and liquid hydrogen (LH2). The baseline two-stage vehicle designated Delta IV Medium has a 4-m-diameter second stage powered by a RL-10B-2 engine using LOx and LH2. Three other configurations of Delta IV Medium vehicles offering progressively more payload weight to low Earth orbit (LEO) or geostationary transfer orbit (GTO) use Alliant Techsystems GEM 60 (60-in. diameter) graphite-epoxy solid propellant motors as strap-on boosters. The Delta IV Heavy uses three of the common 5-m-diameter first stages in parallel. The second stage uses the same longer 5-m-diameter tank used on the Medium+ (5, 2) and (5, 4) vehicles. The numbers in parentheses indicate the diameter of the second stage and payload fairing and the number of graphite epoxy motor (GEM) strap-ons, respectively. This family of vehicles delivers from 20,000 to 48,000 lb to LEO (27.8°), or 9,300 to 28,000 lb to GTO. The propulsion system elements that essentially control performance and risks are the first- and second-stage engines and the solid propellant strap-on motors. These three propulsion systems are summarized in Appendix D. FIGURE 4-2 The Delta IV family of access-to-space vehicles. SOURCE: Knauf (2005). Atlas V Family of Vehicles The Atlas V family of two-stage launch vehicles shown in Figure 4-3 utilizes a 3.8-m-diameter common core first stage. The first stage uses a single rocket engine having dual-thrust chambers (RD-180) operating on LOx and kerosene. The baseline two-stage vehicle designated Atlas V 401 has a 3.05-m-diameter, 12.7-m-long Centaur second stage powered by a RL-10A-4-2 engine using LOx and LH2. The 401 does not use a strap-on solid motor. The various Atlas V configurations available are designated as the 400 Series, the 500 Series, and a heavy lift vehicle (HLV) that was still in the design stage in 2005. The 400 Series has a 4-m-diameter payload fairing and the 500 Series provides a 5-m fairing. The Centaur stage can use either one or two RL-10A-4-2 engines. Depending on the mission, the 500 Series can be configured with from zero to five strap-on solid rocket motors (Aerojet Sacramento). Each motor provides about 254,000 lb thrust at liftoff.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 4-3 Atlas V family of access-to-space vehicles. SOURCE: Knauf (2005). This 500 Series of vehicles can deliver from 20,000 to 45,000 lb to LEO (27.8°) or 8,750 to 19,100 lb to GTO. The propulsion system elements that essentially control the various configuration’s performance and risks are the first- and second-stage engines and the solid propellant strap-on motors. These three propulsion systems are summarized in Appendix D. BOOSTER ENGINES FOR LARGE LAUNCH VEHICLES First-Stage, Liquid Propellant Delta IV: RS-68 In the early 1990s, Rocketdyne initiated development of the first new indigenous booster-class engine in the United States in more than 25 years, the RS-68. The RS-68 was ultimately selected to power the Delta family of EELVs developed for the Air Force by the Boeing Space Systems Company. The RS-68 is the largest LOx/LH2 engine in the world today. It is a conventional bell-nozzle booster engine that develops 650,000 lb of sea-level thrust, the equivalent of 17 million horsepower (or 11 Hoover Dams at full power generation). The engine uses a simple, open gas generator cycle with a regeneratively cooled main chamber. The turbine exhaust gases can be vectored on command to provide roll control. The engine can be throttled to 60 percent of full power. The simplified design philosophy behind this engine meant it had fewer parts and lower production costs than the contemporary space shuttle main engine (SSME). The RS-68 engine has only 11 major components, including the main combustion chamber, single oxygen and hydrogen turbopumps, gimbal bearing, injector, gas generator, heat exchangers, and fuel exhaust duct. This amounts to 80 percent fewer parts than the SSME and a reduction in hand-touched labor of 92 percent. The development cycle time was also much reduced, and nonrecurring costs were claimed to be reduced by a factor of 5 over previous cryogenic engines. The engine was designed, developed, and certified in a little over 5 years and flew on the first Delta IV launch in late 2002. Atlas V: RD-180 Engine The engine that powers the first stage of the Atlas V EELV is the RD-180. The RD-180 is a two-thrust-chamber version of the original Russian RD-170 (four chambers) that is used to power the first stage of the Yuzhnoye/Yuzhmash Ukrainian-manufactured Zenit launch vehicle. This engine provides the

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs performance, operability, and reliability of the RD-170 in a size (933,400 lbf of vacuum thrust) that meets the booster needs of the Atlas V version of the EELV (first used in the United States to successfully power all the Atlas III launches). The RD-180 is an integrated propulsion unit/engine system with hydraulics for control valve actuation and thrust vector gimbaling, pneumatics for valve actuation and system purging, and a thrust frame to distribute loads, all self contained as part of the engine. The engine, which employs a LOx lead start, staged combustion cycle, and an LOx-rich turbine drive, delivers 10 percent better performance than the current operational U.S. booster engines fueled by kerosene rocket propellant-1 (RP-1) and can provide relatively clean, reusable operation (more than one mission duty cycle). Finding 4-2. The current family of U.S. EELV boosters does not need to be replaced for the next 15 to 20 years, nor are there plans to do so. Nevertheless several candidate designs were started under NASA’s Space Launch Initiative (SLI) program in 2001. Recommendation 4-2. DoD should begin work relatively slowly, investing about $5 million per year in the committee’s judgment on technology development for an advanced-cycle booster engine that could provide the basis for a new far-term access-to-space vehicle. First-Stage, Strap-on, Solid Propellant Delta IV+: GEM-60 Alliant Techsystems, Inc. (ATK) originally developed the GEM strap-on solid rocket booster for the Delta II launch vehicle for the Air Force and Boeing. The GEM-40 is a highly reliable motor used on Delta II. The GEM-46 is a larger derivative—with increased length, diameter, and vectorable nozzles on three of the six ground-start motors—for use on the Delta III. The motor has also been used on the Delta II Heavy. The 70-ft GEM-60 motors provide auxiliary liftoff capability (in two or four strap-on motor configurations) for the Delta IV Medium Plus (M+) vehicles. Atlas V: Aerojet The solid rocket strap-on booster motors for the Atlas V were developed, flight qualified, and produced by Aerojet, Sacramento. This new generation of solid rocket motors provides reliable, high-performance boosting power for the Atlas V medium- to heavy-lift expendable launch vehicle used for U.S. civil and military spacecraft launch programs as well as for international and U.S. commercial satellite rockets. The Aerojet solid rocket motor design for the Atlas builds on decades of flight design, test and real mission experience such as the series of Minuteman, Peacekeeper, and small intercontinental ballistic missile (ICBM) motors, as well as Aerojet’s extensive work on other propulsion and space systems and a wealth of accompanying flight proven technologies. The Atlas V family of launch vehicles will use from one to five strap-on solid rocket motors depending upon the mission and launch trajectory requirements. The solid rocket motors are ignited at lift-off and burn for over ninety seconds, each providing a thrust in excess of 250,000 lbf. At about 94 seconds into the flight, the solid rocket boosters are jettisoned sequentially. First Stage, Strap-on, Liquid or Solid Propellant One of the most effective ways to upgrade the payload capability of launch vehicles is to add strap-ons to the first stage. Solid strap-ons have been used frequently for this purpose, but liquids could also be employed. Some of the liquid propellant boosters currently being developed by the various FALCON contractors are of an appropriate thrust level and could be a low cost alternative to solids for this purpose.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs New solid booster technologies under the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program and, possibly, liquid propellant booster concepts that may be developed for a new Air Force responsive spacelift vehicle might be studied for this application. Alternative Hybrid Propellant Strap-ons Lockheed Martin Space Systems has worked on hybrid propulsion technologies since 1989. Its initial studies focused on replacing the solid rocket boosters on the space shuttle after the Challenger disaster. It worked with American Rocket Corporation (AMROC) during the DM-01, DM-02, and Hybrid Technology Options Project (HyTOP) motor development efforts, which eventually led to the Hybrid Propulsion Development Program (HPDP). Within the HPDP, Lockheed Martin tested numerous technologies that were developed under internal independent research and development (IR&D) funding and increased the technology maturity of numerous hybrid-based systems. Under the current FALCON program, it performed a number of tests to demonstrate stable hybrid rocket performance. The largest hybrid motor tested to date using the staged combustion system was the HPDP 250,000 lbf motor, which was approximately 72 in. in diameter and 30 ft long. The tests demonstrated that the system could be successfully scaled to high-thrust motors that could potentially be used for booster or first-stage applications. Second-Stage Engines RL-10 The RL-10 has evolved significantly over the past 42 years. It began in 1963 with a vacuum thrust of approximately 15 klb for the RL-10A-1. Through a series of modifications, the thrust evolved to an average thrust of 24.75 klb in the RL-10B-2. This engine has probably had every last possible ounce of thrust wrung out of it, but that accomplishment has reduced the margins of safety for some of the failure modes. Significant improvements in performance and reliability could be achieved with a new engine-cycle design. Currently, the EELVs have only one basic second stage, the RL-10. The Delta IV uses the RL-10B-2, while the Atlas V uses the RL-10A-4-1 or 2. The basic RL-10 engine, developed in the late 1950s, was the world’s first LOx/LH2-fueled rocket engine operated in space. Since the first successful launch of an Atlas/Centaur RL-10 in November of 1961, Pratt & Whitney has developed nine different models of the RL-10 engine family. The RL-10 had earned the reputation of being a reliable, safe and high-performing cryogenic second-stage engine for a wide variety of upper stages on a large number of U.S. expendable launch vehicles. Current RL10 engine models and their supported vehicles are RL-10A-4-2 (Atlas V), RL-10-4 and RL-10-4-1 (Atlas II, IIA, IIAS, III and IIIB) and RL-10A-3-A (Titan IVB). The full family of flight-certified RL-10-XX engines is listed in Table 4-1, along with the engines’ respective key design features. RL-10A-4-2. The RL-10A-4-2, used on the Centaur IIIB upper stage and the Atlas IIIB and Atlas V launch vehicle, is a LOx/H2 closed expander. It is equipped with a single turbine and gearbox, which drive the two pumps. Additionally, the engine is equipped with dual direct spark ignition and can be flown with a fixed or extendible nozzle. The engine operates nominally with a chamber pressure of 610 psi and develops an Isp of 451 sec.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs RL-10B-2. The RL-10B-2 currently powers the second stage of the Delta III and the medium- and heavy-lift configurations of the Delta IV. It features the world’s largest carbon-carbon extendible nozzle, with an expansion ratio of 285:1. This high-expansion nozzle enables it to operate nominally with a chamber pressure of 633 psi, and develops an Isp of 465.5 sec. TABLE 4-1 Comparison of RL10 Engine Models MODEL NO. A-1 A-3 A-3-1 A-3-3 A-3-3A A-4 A-5 A-4-1 B-2 Vacuum thrust (lb) 15,000 15,000 15,000 15,000 16,500 20,800 14,560 22,300 24,750 Chamber pressure (psia) 300 300 300 395 475 578 485 610 644 Thrust/weight 50 50 50 50 54 67   61   Expansion ratio 40:1 40:1 40:1 57:1 61:1 84:1 4.3:1 84:1 285:1 Specific impulse (sec) 422 427 431 442 444 449 368 451 466.5 Flight certification Nov Jun Sep Oct Nov Dec Aug Feb May date 1961 1962 1964 1966 1981 1990 1992 1994 1998 SOURCES: (1) NASA point paper “Space Propulsion Technology Necessary to Enable Human and Robotic Exploration Missions,” p. 31, R.L. Sackheim et al. (2006), (2) Pratt & Whitney Web site, Pratt-Whitney.com, and (3) Purdue University, liquid rocket engines Web site, Purdue.edu. Finding 4-3. The technology for the RL-10A and RL-10B family of upper-stage engines is now more than 40 years old. Although numerous upgrades have been incorporated over the life of the engine, much of the design is now outdated. Because the second-stage engine for both EELVs comes from a single supplier, Pratt & Whitney, the Air Force is totally dependent on this single contractor and engine for all large payload launches. Should a failure occur that involves the second-stage engine, all launches with these systems would probably be frozen until the root cause is identified and corrected, which could take a year or more. While the probability of such an event is not high, it is not zero. In a time of crisis, this could be extremely debilitating for the nation. The number of failures in recent years (and their cost) would seem to be another good reason for developing and qualifying a new engine that would be supplied by more than one manufacturer. In addition, to make full use of the Delta and Atlas heavy vehicles, a higher thrust engine is needed. To develop a new upper-stage engine for the nation’s fleet of strategic launch vehicles requires a major development effort and an extended qualification program. The extremely high reliability demanded by a strategic launch capability means that a new engine development program may not skimp on hardware or testing. Recommendation 4-3. DoD should place a high priority on development of a new medium-thrust (50,000-80,000 lb) upper-stage LOx/H2 engine to assure the nation’s strategic access to space. The cost of developing such an engine through its initial operation capability (IOC) is estimated by the committee at $150 million to $250 million providing the design does not try to push new technologies to their limits. Alternative Designs for Second-Stage Engines Several government organizations have recommended that a new second stage engine be developed in the thrust class of 50,000 to 100,000 lb. Industry has responded, with Pratt & Whitney developing the RL-60; Rocketdyne the MB-60, and Aerojet the RS-60. Northrop Grumman also has a USET-funded program to design a 40,000 lb LH2 engine. All of these are in various stages of development. The MB-60 has components at TRLs between 6 and 9 depending on the component. Pratt & Whitney teams with several international partners to work on the RL-60. Volvo is producing the nozzle while Ishikawajima-Harima Industries (IHI) is providing the hydrogen turbopump. The RL-60 chamber has been tested.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs Aerojet has worked on the design of its AJ-60 concept but has yet to develop the hardware. It is, however, developing a model that is more heavily physics-based, which helps to mitigate the risk in full engine development, and is advancing the technology for virtual engine design. All of these options offer more thrust than the RL-10 engine, which needs additional capability if heavier payloads are to be placed into higher orbits. New engine design options that appear to be most suitable for the Air Force and DoD missions are compared with the existing RL10A-4 in Figure 4-4. FIGURE 4-4 Upper stage engine options. SOURCES: (1) NASA point paper “Space Propulsion Technology Necessary to Enable Human and Robotic Exploration Missions,” p. 31, R.L. Sackheim et al. (2006), (2) Pratt & Whitney Web site, Pratt-Whitney.com, and (3) Purdue University, liquid rocket engines Web site, Purdue.edu. However, most of these new second-stage engine design efforts are not fully funded. The development of a new rocket engine is a very expensive proposition, costing between $150 and $250 million, providing the design does not try to push new technologies to their limits. Therefore, the large liquid propellant rocket engine industry (now made up of only three companies) have not been able to justify committing large amounts of increasingly scarce internal resources to full development and qualification of any of these candidate concepts. Support by industry could increase differently if DoD were to commit to a serious, well-funded, long-term program for a new family of large, responsive spacelift vehicles to support major new total capability in-space architecture. SMALL TO MEDIUM-SIZED LAUNCH VEHICLES Existing Vehicles Pegasus The Pegasus is a vehicle launched in midair (via a modified Lockheed L-101 I aircraft) (OSC, 2000). Orbital Sciences Corporation (OSC) manufactures the three-stage, all-solid-propellant, three-axis stabilized vehicle. The Pegasus-XL vehicle, a stretched version of the original Pegasus vehicle, can place a 400- to 1,000-lb payload into low Earth orbit. The original version of the Pegasus was retired in 2000, and only the Pegasus-XL is used today. The Pegasus-XL free falls for 4 seconds after release; then the first-stage solid rocket motor, manufactured by Alliant TechSystems, fires and burns. The delta-shaped wing produces lift, and the launch vehicle begins a 2.5 g force pull-up. Then the second-stage solid fuel motor ignites, and at approximately 2 minutes, the payload fairing is ejected. The second stage flies to an altitude of approximately 129 miles with a velocity of over 12,000 miles per hour. At the appropriate altitude to achieve the designated orbit, the third-stage motor ignites and burns for 1 minute and 6 seconds to place its payload into orbit.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs NASA certified Pegasus to carry the highest value satellites (Category 3 certification) because of its excellent reliability record. Pegasus has launched its last 21 missions successfully. No Pegasus XL vehicles flew in 2004. On April 15, 2005, a Pegasus XL successfully launched the demonstration of autonomous rendezvous technology flight demonstrator vehicle for NASA (FAA, 2006). Athena The Athena I carries a payload of up to 1,750 lb and the Athena II, up to 4,350 lb.3 The Athena I and II use Thiokol’s Castor 120 motor with 435,000 lb thrust for their first and first and second stages, respectively. The engine burns hydroxyl-terminated polybutadiene (HTPB, a polymer) propellant. The Athena I second stage and the Athena II third stage are powered by Pratt & Whitney’s Orbus 21D, with a thrust of 43,723 lb. Athena I and II have a common orbit adjust module that houses the attitude control system and the avionics subsystem. The monopropellant hydrazine fuel (a liquid) attitude control system performs orbital injection corrections, roll control, velocity trim, and orbit circularizing maneuvers. The first operational mission of the Athena, an Athena I, successfully launched the NASA Lewis satellite into orbit from Vandenberg Air Force Base in California, on August 22, 1997. The first Athena II was successfully launched from Cape Canaveral, in Florida, on January 6, 1998, sending NASA's Lunar Prospector spacecraft on its mission to study the moon. Subsequent successful missions include Athena I/ROCSAT-1 for the Republic of China on January 26, 1999, Athena/IKONOS for space imaging from Cape Canaveral on September 24, 1999, and Athena/Kodiak Star for NASA from Kodiak, Alaska, September 29, 2001. The Athena I and II use a simple and reliable orbit adjust module (OAM) that can be adapted to many small launch vehicle configurations. The OAM houses the attitude control system and avionics subsystem (guidance and navigation, batteries, telemetry transmitters, command and destruct receivers and antennas) that are common to Athena I and Athena II. The OAM is located directly beneath the payload to perform the final orbit injection burns and any needed to put the satellite in the precise orbit. The OAM weighs 819 lb dry and can carry no more than 960 lb of hydrazine. After payload separation, the OAM performs a contamination and collision avoidance maneuver, distancing itself from the payload and burning any remaining fuel to depletion. The attitude control system, provided by Aerojet, uses off-the-shelf propulsion components. The propellant load is tailored to the specific mission. Taurus Taurus is a ground-launched version of the OSC’s Pegasus rocket vehicle. It uses three stages of the Pegasus boosted by a large Castor solid propellant motor. It is designed to launch satellites up to 3,500 lb into LEO. Liftoff weight varies between 150,000 and 220,000 lb. It can be transported and launched from various minimally improved sites in the world.4 Four variants of the Taurus launch vehicle exist. The smallest version, known as the ARPA Taurus, uses a Peacekeeper first stage instead of a Castor 120 motor. A second size uses the C-120 first stage and a slightly larger Orion 50S-G second stage. The Taurus XL uses the Pegasus XL rocket motors (Orion 50S-XL and Orion 50XL) and is considered a development-stage launch vehicle. The largest Taurus variant, the Taurus XLS, is a study-phase vehicle that adds two Castor IVB solid rocket boosters to the Taurus XL to improve payload by 40 percent over the standard Taurus. For all Taurus configurations, satellite delivery to a GTO orbit can be achieved with the addition of a Star 37FM perigee kick motor. 3 For additional information, see http://www.lockheedmartin.com/wms/findPage.do?dsp=fec&ci=11459&rsbci=0&fti=0&ti=0&sc=400. Last accessed on March 30, 2006. 4 For additional information, see the Taurus fact sheet at http://www.orbital.com/NewsInfo/Publications/Taurus_fact.pdf. Last accessed on November 19, 2006.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs The Taurus system is evolving more responsive payload integration and launch operations. Second stages are integrated horizontally and payloads are integrated with the fairing in a separate area. This format for operations will be an almost mandatory part of the total system architecture of future operationally responsive launch systems. Minotaur For the Air Force's Orbital/Suborbital Program (OSP), Orbital developed the low-cost, four-stage small launch vehicle (SLV) Minotaur rocket using a combination of U.S. government-supplied Minuteman II motors as the vehicle’s first and second stages and proven OSC space launch technologies (OSC, 2004). Minotaur's third and fourth stages, structures, and payload fairing are common with the Pegasus XL rocket. Its capabilities have been enhanced by adding improved avionics systems, including a modular avionics control hardware, which is used on many of OSC’s suborbital launch vehicles. Minotaur is considered a small launch vehicle. It can lift 750 lb to a 400-nm, sun-synchronous orbit. This is roughly 1.5 times the Pegasus XL capability. All payload customers must be U.S. government agencies or be sponsored by such agencies. The Secretary of Defense holds approval power for each launch mission. Sea Launch In 1995 the Sea Launch Company, LLC, headquartered in Long Beach, California, was formed, with 40 percent owned by Boeing, 20 percent Kaevener (Norway), 25 percent Energia (Russia), and 15 percent Yuzhnoye/Yuzmash (Ukraine). The Sea Launch system combines launch, home port, and marine segments to offer a heavy-lift capability of 6,000 kg and injection into GTO from a performance-enhancing equatorial launch site. The launch segment consists of the Zenit-3SL rocket produced by Yuzhnoye/Yuzmash in Dnepropetrovsk, Ukraine; the Block DM-SL upper stage produced by Energia in Moscow; and payload accommodations produced by Boeing in Seattle. The first and second stages of the Zenit-3SL are powered by the RD-171 and RD 120 LOx/kerosene engines, respectively. The Block DM-SL upper stage is powered by the 11D58M LOx/kerosene engine. The payload accommodation module consists of a graphite epoxy 4-m diameter payload fairing and a payload interface adapter. All launch vehicle processing, spacecraft processing, and payload encapsulation takes place at home port in Long Beach. Payload processing is managed by Astrotech in Sea Launch’s payload processing facility. The marine segment consists of the launch platform (LP) and the assembly and command ship (ACS). The ACS encompasses the launch control center, range safety, a weather station, and accommodations for crew and customers. Operationally, the Zenit-3SL is integrated horizontally within the ACS then transferred to the LP. During the trip to the maritime equatorial launch site at 154°­­­ W (11 days for the LP and 8 for the ACS), launch operations and rehearsals are conducted to ensure crew readiness. On April 26, 2005, Sea Launch successfully delivered a Boeing 702 model spacecraft weighing 6,080 kg into GTO. The Boeing Sea Launch Web site indicates there were 22 successful launches from this system through June 2006.5 Finding 4-4. The Sea Launch operations concept could provide key advantages for a variety of small to medium-size launch vehicles and needs to be seriously considered as a viable launch vehicle for military geosynchronous payloads even though ownership is multinational. For GTO missions, launches are conducted from a point on the equator at approximately 154° W. This has two significant performance advantages. It allows Zenit to deliver spacecraft to a GTO transfer orbit at roughly 0° inclination, thus reducing the required spacecraft apogee burn and allowing the Zenit 3SL rocket to lift a heavier spacecraft mass or provide longer life in orbit. The maritime launch also nearly eliminates range safety 5 For additional information, see the official Boeing Sea Launch Web site at http://www.boeing.com/special/sealaunch/. Last accessed on August 8, 2006.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs be selected early on, since it will probably have the most influence on the operational factors involved in total mission success. Upper-Stage Engines USET is an important step in achieving the computerized, high-fidelity virtual engine designs that can be tested on a virtual test stand. It is anticipated that close correspondence between predicted performance and delivered performance, especially under transient and highly off-nominal run conditions can become a reality. However, achieving this goal is highly unlikely given the funding and schedule constraints of the current USET contract. Because this program is one of the key elements of a total architecture systems engineering process and can have a huge impact on the development schedule and cost of ARES, it needs to be realistically supported in the immediate future. Reusable Booster Engines To develop a first-stage reusable engine for the ARES subsystem demonstrator, it would be extremely useful to take all of the lessons learned during the development, certification, and upgrades of the SSME, which is the only flight-demonstrated, reusable booster rocket engine in the world, and see what can be applied to ARES booster engine needs. Committee members who visited the Rocketdyne site reported that this had been done in great detail for the Delta IV RS-68 engine and the RS-83 and RS-84 next generation of reusable engines being designed for the NASA SLI and NGLT programs. Pratt & Whitney and Aerojet reported the same approach for their COBRA design for the same program. Advanced Pumps, Turbines, and Power Cycles Other than the IPD currently being tested under NASA funding at Stennis, little new engine design or development work is being funded by DoD. Much effort is needed to validate design criteria for advanced pumps, turbines, new power cycles such as oxygen-rich staged combustion, single preburners, high-pressure (greater than the RL-10), high-performance expander cycles, and other engine system elements. The validation of technologies such as these will be crucial for the success of new rocket engines that can meet the performance and risk objectives of ARES and future ORS. New Propellants Energetic yet insensitive propellants will be needed to develop low-cost, high-energy propellants for use in both solid propellant motors and liquid propellant engines. Prospects for the introduction of very high bond energy fuels in the near term seem doubtful. Higher density, higher Isp monopropellants may evolve first, but even if one of them is validated it could take many years to establish a reliable industrial facility for producing it at an acceptable cost. The most significant problem facing future higher energy density systems is developing appropriate materials, because nearly all such fuels will have to operate at higher chamber and nozzle temperatures to produce even equivalent Isp owing to the higher molecular weight of their combustion products. Nevertheless, a strong, continuous effort is required just to come up with options for the far term. Effects of Solid Propellant on Motor Aging Better technologies for measuring aging need to be developed for solid rocket motors. Surveillance techniques are required so that individual motors that have aged out can be identified and removed from inventories.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs Storable Propellants Several storable oxidizers and fuels that have been used by the United States for both launch vehicles and in-space propulsion systems for more than 55 years are toxic. Because of this, they are considered difficult and expensive to handle safely for certain types of missions (primarily manned and civilian-operated launches). New less toxic oxidizers and fuels that can be stored for indefinite periods both on the ground and in space could enhance the chances of mission success. Such propellants should have nearly the same performance—e.g., Isp, density Isp, safety, handling, and operation—as current combinations. For some applications, high specific density impulse may offset the higher specific impulse of nonstorables. New Hydrocarbon-Fueled Rockets Rocket engines using advanced hydrocarbons will require stable operation. Current stability models are still of only limited usefulness for real engine design. More physics-based models are needed. It would be cost prohibitive now to carry out hundreds of tests such as were carried out for the development of the F-1 engine. In addition, the hydrocarbon fuels need to be thoroughly characterized because variables can have critical impacts on stability limits. Researchers need to do trade studies of the impact on performance of using different fuels, including fuels with different density and energy contents. Managers must also consider the cost of the infrastructure for making a particular fuel available to a launch site vs. the loss of performance of using other more widely available fuels. Ablation Rates Better characterization and optimization of ablation rates for different materials are needed. The Office of Naval Research has a program under way with which some synergy might be possible, but the thrust levels involved in that program are much smaller than those of SLVs. An urgent problem to be solved is finding the transfer functions for the rates as a function of pressure, mixture ratio, geometry, and so on. For ablative nozzles, both the selection of materials and the manufacturing process are important for the final performance of the nozzles. Continuing work on technology in both areas will be required to accommodate new propellant characteristics. Applying Lessons from SSME to New Designs The SSME is the only mission-demonstrated reusable booster rocket engine in the world. The information gained from that experience base regarding technology limits, failure modes, manufacturing issues, ability to control the engine configuration after high-stress reuse, and other problems needs to be explicitly applied to the conceptual design phase of ARES reusable booster engine candidates. Of course, many of the lessons learned from SSME apply mainly to the LOx/LH2 propellant combination. As described earlier, there are three large existing engines with potential for reuse as boosters plus four engines that can be scaled up and have some potential for reuse. Several totally new engine concepts where designs have different levels of maturity and that incorporate technologies using different propellants and having different levels of design-criteria validation were also discussed. Recommendation 4-11. The Air Force should develop two reusable liquid propellant first-stage rocket engines for the ARES demonstrator launch vehicle. The design of these engines should take advantage of all the engineering lessons learned during the development, certification, and extensive upgrades of the SSME. To permit the Air Force to have dual-source propulsion systems for ARES and subsequent ORS vehicles, two engine design concepts should be selected based on different propellants and configurations having functional and hardware failure modes as different as possible.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs AREAS THAT DESERVE MORE ATTENTION Physical and Thermodynamic Properties of Fuels and Oxidizers In 2004, the Joint Army, Navy, NASA, Air Force (JANNAF) Liquid Propulsion Systems Committee formed a panel on hydrocarbon fuels to investigate current models for RP-1 properties and develop modern specifications for heat of combustion, viscosity, density, and thermal stability. That committee recommended that thermal stability testing include the jet fuel thermal oxidation tester method traditionally used for jet fuels and the high Reynolds number thermal stability (HiReTS) method. The HiReTS method had never before been applied to RP-1. There are only two operational HiReTS testers, one at the Southwest Research Institute in San Antonio, Texas, and the other at the UAH. UAH is performing HiReTS tests on traditional RP-1 with low sulfur and various additives (red dye). NASA Glenn Research Center has an ongoing program for characterizing the thermal response of hydrocarbon fuels. This effort includes use of a heated tube facility to study thermal stability, coking, and the heat transfer properties of jet fuels and RP-1. Trade studies are needed on the performance impact of using different fuels to meet specific missions, including variability in properties such as density and energy content. Storable Oxidizers Updating the physical and thermodynamic properties of oxidizers that are indefinitely storable in space using practical thermal control systems (or that could be made storable) is considered an enabling technology for many on-orbit applications. With the launch of the last storable-oxidizer Titan IV in 2005, all remaining major U.S. liquid-propellant launch vehicles use LOx for the oxidizer. This cryogenic fluid requires special high-maintenance storage and handling equipment along with a large team of engineers and trained technicians at every launch site. These issues were a primary reason that the Thor and Atlas and Titan I ballistic missiles were phased out in favor of the Titan II, which used storable oxidizer, and Minuteman-type missiles, which used solid propellant. For the same reasons, the cost and risks of achieving full-time readiness to fuel and launch ORS systems at a large number of sites worldwide will be very high if the vehicles are committed to using LOx. A number of storable oxidizers are candidates for ORS in the optimization of mission-success-based total systems engineering. For example, the reasons NASA does not like N2O4 for reusable shuttle operations are not compelling for ORS-type military missions. An alternative to pure N2O4 is a 65/35 mixture of N2O4/N2O. This storable oxidizer with RP fuels can produce a quite high specific density impulse. The adaptation of high-energy, storable oxidizers to boosters and upper stages could be one of the few technology areas that might have the potential to engender a transformative storable system for rocket-propelled access to space or near space. One approach to consider is the encapsulation by means of nanotechnology of high-energy, non-liquid oxidizer molecules stable-slurried in medium-performance liquid oxidizers such as H2O2 or N2O4. Examination of the work in nanotechnology energetics for explosives and monopropellants at Pennsylvania State University and the University of Southern California could be starting points for an R&T program focused on high-energy, storable liquid oxidizers. Finding 4-12. A program is needed to explore various approaches to creating storable oxidizers that would provide significantly increased rocket performance with different storable fuels. Recommendation 4-12. DoD and the Air Force should fund a program to explore various approaches to creating storable oxidizers that would significantly enhance rocket performance with different storable fuels. This program should utilize a consortium of academic, industry, and government laboratories to pursue highly innovative concepts for achieving this breakthrough.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs Materials Input on materials issues as they affect rocket propulsion was received from all of the contractors visited and from the Materials and Manufacturing Directorate of the AFRL. Major efforts are ongoing at Aerojet, Rocketdyne, and Pratt & Whitney, with less ambitious efforts at other contractors. Materials are frequently a pacing factor in the development of advanced propulsion systems, and rocket engines are no exception. On the LOx side, improved high-strength nickel-based alloys are in the early stages of development, but much more testing and validation is required. Nanocrystalline aluminum alloys also show promise for pump components and lines. On the fuel (hydrocarbon) side, improved titanium alloys could be developed for pump components. Unlike aircraft gas turbines, the turbine section of the turbopump operates at relatively low temperatures but much higher pressures. This is because the turbine is operating either fuel-rich or LOx-rich. The higher pressures produce operating conditions much different from those in a conventional aircraft gas turbine. Thermal barrier coatings for copper alloy combustion chamber walls could reduce thermal strains and, in addition, mitigate coking issues. This is an area of research that would be particularly well suited to a consortium approach, as mentioned earlier. Availability of high-conductivity copper alloys (NASA-Z for example) is also a continuing issue and needs to be addressed. For SRMs, organic matrix composites have been developed that have resulted in significant performance improvements. However, fiber availability continues to be a problem, and rayon for nozzle applications is now sourced overseas. Improved nondestructive inspection systems are in various stages of development, but validation of these approaches to avoid the destructive testing of aging SRMs is a priority. For in-space propulsion, development of oxidation-resistant materials is a priority. In particular, materials having the performance of Ir/Re alloys but at a lower cost are needed. Improved high-temperature insulation materials are also needed. Ceramic matrix composites (CMCs) are a potential candidate for this application and are also being considered for cooled combustion chambers. Finding 4-13. All of these materials requirements for in-space propulsion need to be balanced against the changing and maturing Air Force and DoD needs and then adequately funded to assure a TRL level of 6 or higher by 2018. Recommendation 4-13. A consortium of industrial partners and the government would appear to be the optimum solution in several of these areas and was demonstrated to be effective for the development of turbine engine materials and processes. Propulsion Element Technologies Turbopumps A turbopump is one of the most highly stressed components of a rocket engine and therefore one of the most trouble prone. Bearings and seals operate in a relatively hostile environment and at very high speeds, with rapidly changing load transients. There are issues with rotordynamic instability, fatigue, oxidation, hydrogen embrittlement, and cavitation. Many of these problems are addressed analytically with existing tools with varying degrees of uncertainty. However, extensive testing is still required in most cases to establish design criteria, performance spreads, and failure mode margins, particularly if multiple reuse is contemplated. As virtual engine capabilities evolve, much of the expensive and time-consuming early cut-and-dried testing of turbopump components can be eliminated. Significant testing will still be necessary, but it will be focused on final qualification of flight hardware and establishing risk uncertainty profiles. Eventually, that test data bank will upgrade the virtual engine design capability so

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs that narrowing the risk uncertainty profile of final flight engines (integrated with their propulsion systems) can be done with minimal losses of expensive hardware at greatly reduced test program duration and cost. The Air Force USET team has chosen a variety of software from various subcontractors as the basis for turbopump analysis and design. There are two USET contracts, one with Northrop Grumman and one with Aerojet, both with the same goals: to develop models for turbopump analysis and design. In addition, the skills required to design a high-performance turbopump are very specialized and must be learned on the job. Critical skills retention is a major issue for the nation if the capability to design future rocket launch systems is to remain world-class. Hybrid Technology Insulation materials compatible with hybrid combustion products need to be improved to accomplish run-to-empty (i.e., no residual fuel) operation. Future hybrid motor insulators need to serve as a structural element during the initial burn, when the chamber pressure loads are highest and to withstand erosion when exposed. Materials testing in a relevant environment will allow minimizing fuel residuals and decrease the inert mass devoted to insulation. Test data indicate that nozzle throat materials typically used for solid propulsion systems, such as three-dimensional carbon-carbon and ATJ graphite, erode relatively quickly in a high-pressure hybrid combustion environment. Significant erosion of the nozzle throat does not affect the hybrid fuel burn rate, but it does reduce the nozzle expansion ratio and chamber pressure as a function of time, which eventually degrades performance. Either future nozzle materials that are more compatible with hybrid propulsion need to be identified and developed or cooling techniques, such as film cooling with fuel or oxidizer, need to be employed to reduce throat erosion rates well below 5 mil/sec for long-duration motor burns. Reliability of the Supply Base The committee visited more than 10 contractors serving DoD and NASA as suppliers of rocket engines and rocket engine components. All of the contractors expressed concern about the viability of the supply base, particularly in the area of specialized materials. Some of the specific supply base issues of concern to contractors pertained to the following materials: High-conductivity copper base alloys for combustion chambers, Advanced titanium alloys, Rayon fiber for composite rocket motor casings, and Ceramic matrix composites. Many of these materials are required only for rocket engine applications, resulting in limited and disjointed demand. Finding 4-14. Advanced materials are required for the continued development of high-performance rocket propulsion systems, and certain of these materials have specialized uses in rocket engine applications. Availability of these advanced, but specialized, materials will be key to the success of future space initiatives. The cost of developing and qualifying some of these new materials and maintaining qualified suppliers could probably be reduced by forming appropriate industry-government consortia. Recommendation 4-14. DoD and the Air Force should take the lead in establishing viable methods to achieve availability and assured continuous supplies of critical materials and items, including new ablative materials for thermal insulation and new materials for ITE nozzles for high-temperature and high-pressure applications.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs LEVERAGING OPPORTUNITIES FOR ACCESS-TO- SPACE PROPULSION Leveraging resources, data, and technology development available from DARPA, NASA, industry, and academia could reduce the time and costs to the Air Force of developing certain technologies for smallsat launch vehicles, ARES, and future ORS systems. Low-Cost, Responsive Launch Vehicles The DARPA FALCON program goals were aimed at revolutionizing the way the United States designs and builds launch vehicles so that they would have more aircraft-like operations while still being cost-effective. In September 2005, DARPA downselected to AirLaunch for Phase IIB. Perhaps some of the propulsion system modules to be demonstrated could be modified, combined, and/or scaled to leverage the development of propulsion systems for ARES and, eventually, for some of the ORS vehicles. Some of the AirLaunch data might be leveraged to support other air launch concepts such as airborne vertical launch and multimission modular vehicles. Other small, expendable launch vehicles being developed by industry—for example, SpaceX—could also provide an opportunity for supplying technologies for small access-to-space vehicles for the Air Force. Propulsion Technologies Developed by NASA The current launch vehicle architecture being evolved by NASA does not offer many propulsion elements for direct leveraging. However NASA’s fundamental R&D programs and extensive library of design criteria should offer many opportunities to leverage systems engineering and vehicle design tools, physics modeling, new materials, and other technologies. If more industry groups would commit to ARES as a start toward new access to space, they could end up by developing and launching small affordable payloads and satellites. This in turn could stimulate the market and increase the U.S. annual launch rate. The overall cost of producing and launching expendable vehicles would be reduced through economies of scale from the increased production of more of the same basic vehicles. Affordable access might rejuvenate the U.S. aerospace market and reverse the current decline. A reinvigorated interest in launch vehicle design and development would increase competition within the industry and lead to more employment opportunities for young aerospace and propulsion engineers. This in turn would evolve synergies and capabilities that would present future leveraging opportunities for the Air Force in evolving its access-to-space total architecture. STATUS AND CAPABILITIES OF THE U.S. ROCKET PROPULSION INDUSTRY The U.S. rocket propulsion industry and associated space transportation business have been in a steady state of decline since the end of the Apollo and the ICBM cold war missile race era (circa 1972). A turnaround in the propulsion and space transportation industry was expected after the space shuttle and, subsequently, the International Space Station programs were authorized to proceed. The shuttle program of the National Space Transportation System, which had to develop three new liquid rocket engines—the SSME, the orbital maneuvering engine, and the RCE—and the world’s first large, segmented, reusable-case solid rocket motor, did not reverse the decline from the Apollo era; it only slowed the rate of decline until the late 1970s (see Tables 4-10 through 4-14).

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs TABLE 4-10 Historical Trends in National Rocket Propulsion Funding as a Percentage of Apollo Program Peak Funding by Year Year Share of Peak Funding(%) Comments 1967 100 Apollo peak propulsion effort. 1970 80 Beginning of post-Apollo slow down (after Apollo-11, first lunar landing). 1972 20 Apollo essentially finished. Shuttle or NSTS era begins. 1982 17 Propulsion for shuttle does not turn around the rate of propulsion R&D decline but merely slows down the rate of decline. 1986 35 2 year funding spike from Strategic Defense Initiative (Star Wars) funding infusion. 1990 15 Funding profile bottoms out. 1997 15 Propulsion R&D funding stays flat at minimal levels. 2001 23 New NASA R&D propulsion initiatives for SLI and NGLT Pathfinder, X-33, X-37, etc. are funded for a while. 2005 10 All NASA propulsion initiatives are terminated except IPRHPT, minimal propulsion R&D funding for U.S. government. SOURCE: Sackheim (2006). TABLE 4-11 U.S. Rocket Engine Developments from 1955 to 2005 Period Number of New U.S. Engines Developed and Flown Comments 1955-1962 9 Perceived ICBM gap missile crisis with USSR post-Sputnik response. 1963-1967 14 Continued buildup of U.S. ICBM missile capabilities. Apollo race to the moon against USSR. 1968-1972 2 Post-Apollo slowdown. Standardization of SLVs. 1973-1984 6 Space shuttle propulsion development, some SLV upgrades. 1985-1987 3 Shuttle upper-stage motor, PAM-A, PAM-D, etc., but all solids. 1988-2001 0 No new U.S. launch vehicle engines, rest of world develops 40-50 new engines. 2002-2005 1 RS-68 for Delta IV. SOURCE: Sackheim (2006).

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs TABLE 4-12 Cancelled Propulsion Programs TABLE 4-13 NASA and Support Contractor Employment 1960-2000 Period/Year Number of Employees, Average over the Period Number of NASA Employees over the Period Program/Era 1960 120,000 20,000 Apollo 1967 395,000 30,000 Apollo 1972-1985 130,000 35,000 Post-Apollo shuttle, etc. 1987-1995 230,000 20,000 International Space Station 1996-2002 180,000 20,000 Fits and starts: SLI, NGLT, etc. SOURCE: Sackheim (2006). TABLE 4-14 Total NASA Space Transportation Budget 1959-2000 (billion FY97 dollars) Period/Year Budget Comments 1959 2.0 Start of NASA 1965 14.5 Apollo peak 1970 2.5 Apollo roll-off 1974-1982 5.5 Shuttle 1982-1990 6.5 Shuttle continuation plus International Space Station 1994 5.2 Shuttle sustaining 2000 4.0 Shuttle plus some ELVs SOURCE: Sackheim (2006). In the United States, the development of technology for rocket propulsion, for all spaceflight applications has significantly lagged behind that in the rest of the world since the initial certification of the space shuttle. This lack of progress in advancing rocket propulsion technologies over such a long period has resulted in several deficiencies in today’s U.S. national space program. Most notable is the reduced reliability of U.S. launch and space vehicles, as evidenced by the increased number of flight

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs failures during the late 1990s and into this new decade, as well as by the country’s shrinking share of the global market in both the space launch and spacecraft industries. The U.S. launch market share fell from about 80 percent in the late 1970s to less than 20 percent worldwide in 2002 (Sackheim, 2006). Again, in 2005, the U.S. market share was only about 25 percent of the launches conducted worldwide (~15 out of 58). Of these, about half the U.S. launches used Russian engines (i.e., RD-180) and major Russian/former Soviet Union components, and in some cases, complete vehicles (e.g., Zenit for Boeing Sea Launch and Protons for Lockheed Martin/ILS Krunischev). In the last three decades, only one new U.S. government-sponsored booster engine, the SSME, has been developed and gone through flight certification. Some significant upgrades have been incorporated into the SSME since its original certification for flight in the 1970s. These upgrades increased reliability and safety and somewhat increased mean time between engine refurbishment. They did not appreciably advance rocket engine technology. Since 1970, the number of firms capable of major engine development has shrunk significantly. This industry downsizing, combined with consolidation, points up the diminution of the nation’s ability to meet DoD’s propulsion needs for a new ORS family of vehicles starting with ARES around 2015. Basically, our current capabilities in space propulsion and space transportation are but a fraction of the capabilities we amassed starting in 1954 with the ICBM programs and culminating about 1970 with the end of Apollo programs. These programs helped the United States to respond to international crises and to eventually win the cold war. Since 1980 only one new first-stage rocket engine has been developed in the United States. This engine, the RS-68, was funded primarily by Boeing Rocketdyne Propulsion and Power. It was developed as a low-cost expendable booster engine for the Delta IV EELV. Engine performance of the RS-68 is poorer than that of the 1960s-era Saturn V second- and third-stage J-2 engines, both of which were simple open-cycle, gas-generator-powered designs. However, important advancements in engineering methodology and capability were made by the developer through incorporation of comprehensive modeling, computer-aided design/manufacturing, and advanced manufacturing technologies of the 21st century. This later manufacturing technology would be very beneficial in production runs of, say, 30 to 50 engines per year. However, it turns out the EELV program will require no more than five to eight engines a year. The near-term commercial space marketplace for very large boosters has not materialized. As a result, the RS-68 offers almost no unit cost advantage over older engines that are available in several places in the world today. While the United States has developed almost no new booster rocket technology during the last 30 years or more, the new spacefaring nations of Europe, Asia (including India), and the Middle East have been developing their own new vehicle and propulsion systems to catch up. They, along with the former Soviet Union, are believed to have developed 40 to 50 new engines using several propellant combinations in addition to LOx/LH2. Many of these engines can now be considered to be today’s state of the art. Based on these observations, it is probably no coincidence that both the total U.S. share of the space launch market and the reliability of U.S.-built launch vehicles have eroded badly in the last 40 years. In the commercial space marketplace alone, the United States now captures only about $1 billion to $2 billion out of a potential worldwide commercial launch market of $8 billion to $10 billion per year. A similar trend has been observed in development of the upper-stage and in-space propulsion technology products. Advancements in both will be crucial for future Air Force capabilities. Most of the U.S. in-space propulsion developments in recent times have been privately funded with some support from the government. Even here, most of the government-sponsored projects were stopped for one reason or another before any significant advances in technology readiness could be achieved. Finding 4-15. The severe industry downsizing and consolidation causes concern about U.S. ability to meet the propulsion needs set forth in the SMP FY06 (AFSPC, 2003) for a new operationally responsive family of spacelift vehicles, starting with ARES in 2010 and ORS in 2015. DoD and Air Force commitment to fully develop these new robust launch vehicles might help rejuvenate the U.S. aerospace industry, provide more employment opportunities for young aerospace engineers, and reverse the current

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs decline in rocket propulsion design, development, testing, and production capabilities. This in turn could create synergies and capabilities that would present future leveraging opportunities for the Air Force. Recommendation 4-15. The Air Force and DoD should devote more of the annual S&T rocket propulsion budget resources over the next few years to rocket propulsion technologies that would enable the successful introduction of mission-based ORS, and to other flexible, small-satellite launch capabilities in the medium term. The committee’s estimate of the additional focused investments needed is $50 million to $75 million annually. REFERENCES Published AFSPC (Air Force Space Command). 2003. Strategic Master Plan FY06 and Beyond. Peterson Air Force Base, Colorado. Available online at http://www.peterson.af.mil/hqafspc/library/AFSPCPAOffice/Final%2006%20SMP--Signed!v1.pdf. Last accessed on March 30, 2006. AIAA. 1998. Guide for Verification and Validation of Computational Fluid Dynamics Simulations, AIAA-G-077-1998. Blackmon, James B., and D. Eddleman. 2005. Reciprocating Feed System as an Alternative to Turbopump and Conventional Tank Pressurization Propulsion Systems. Propulsion Research Center White Paper, Department of Mechanical and Aerospace Engineering, University of Alabama in Huntsville. March. DARPA (Defense Advanced Research Projects Agency). 2004. FALCON Force Application and Launch from CONUS: Small Launch Vehicle (SLV) Phase II. Program Solicitation Number 04-05 Task I. May 7. Available online at http://www.darpa.mil/tto/solicit/falcon_ph2slv.pdf. Last accessed on March 30, 2006. DoD (Department of Defense). 2004. Department of Defense Space Science and Technology Strategy. Washington, D.C.: Defense Research and Engineering. July 31. FAA (Federal Aviation Administration). 2006. U.S. Commercial Space Transportation Developments and Concepts: Vehicles, Technologies, and Spaceports. Office of Commercial Space Transportation. January. Available online at http://ast.faa.gov/files/pdf/newtech2006.pdf. Last accessed on March 30, 2006. Hampsten, Ken, Jim Ceney, and Gus Hernandez. 2005. ARES Subscale Demo Phase I. Presentation to ARES Industry Day, El Segundo, Calif. March 7. James, Larry. 2005. ARES Industry Day Welcome and Introductions. Presentation to ARES Industry Day, El Segundo, Calif. March 7. Joyner, C. Russell, and Patrick M. McGinnis. 2004. The Application of ITAPS for Evaluation of Propulsion and Power at the System Level. AIAA Paper 2004-3849. Knauf, Jim. 2005. DoD Space Transportation Perspective. Presentation to the First Meeting of the NASA Exploration Transportation Strategic Roadmap Federal Advisory Committee, Orlando, Fla., February 3-4. Available online at http://www.hq.nasa.gov/office/apio/pdf/cev/11_dod_perspective.pdf. Last accessed on March 30, 2006. NRC (National Research Council). 2004. Evaluation of the National Aerospace Initiative. Washington, D.C.: The National Academies Press. Available online at http://www.nap.edu/catalog/10980.html. Last accessed on March 31, 2006. NSPD (National Security Presidential Directive). 2005. U.S. Space Transportation Policy. NSPD-40. January. OSC (Orbital Sciences Corporation). 2000. Pegasus Users Guide. August. Available online at http://www.orbital.com/NewsInfo/Publications/peg-user-guide.pdf. Last accessed on March 30, 2006.

OCR for page 108
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs OSC. 2004. Minotaur Users Guide. October. Available online at http://www.orbital.com/NewsInfo/Publications/Minotaur_Guide.pdf. Last accessed on March 30, 2006. Sackheim, Robert. 2006. The status of propulsion in the U.S. today. Journal of Propulsion and Power. November. Shelton, J.D., R.A. Frederick, Jr., and A.W. Wilhite. 2005. Launch Vehicle Propulsion Design with Multiple Selection Criteria. AIAA Paper 2005-3581. July. Weeks, David, Steven H. Walker, and Robert L. Sackheim. 2005. Small satellites and the DARPA/Air Force FALCON program. Acta Astronautica: AA2399. Wood, Byron. 2002. Propulsion for the 21st Century—RS-68. AIAA Paper 2002-4324. Yang, Vigor, and William E. Anderson, eds. 1995. Liquid Rocket Engine Combustion Instability. Washington, D.C.: American Institute of Aeronautics and Astronautics. December. Yang, Vigor, Mohammed Habiballah, James Hulka, and Michael Popp, eds. 2003. Liquid Rocket Thrust Chambers: Aspects of Modeling, Analysis and Design. Reston, Va.: American Institute of Aeronautics and Astronautics. February. Unpublished Roy Hilton. “Vehicle system M&S and assessments,” Presentation to the committee on April 14, 2005. Mike Huggins. “IHPRPT overview,” Background information provided to the committee on April 14, 2005. Joe Leahy. “P-STAR: A propulsion sizing, thermal analysis, accountability, and weight relationship first-order modeling tool,” Background information provided to the committee on March 30, 2005. Ron Sega. “National Aerospace Initiative update,” Presentation to the committee on March 1, 2005. Scott Smith. “Air-based vertical launch,” Presentation to committee member Gerard Elverum on December 19, 2005.