National Academies Press: OpenBook

A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006)

Chapter: 4 Rocket Propulsion Systems for Access to Space

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Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

concerns and the need to shape trajectory as no populated areas are near the launch site or the downrange impact areas.


Recommendation 4-4. DoD should incorporate this concept into some of the total systems architectures options to be studied for future operationally responsive access to space.

Vehicle in Development

Kistler K-1

The Kistler K-1 vehicle is a two-stage fully reusable vehicle that is being designed for 100 flights, a 9-day turnaround, and 3-day response. Both stages return to the launch site for refurbishment and reuse and use horizontal vehicle processing and checkout.

The K-1 uses the NK-33 engine that was developed by the Russians and designed for multiple starts with large margins for robustness. The K-1 vehicle uses three NK-33s in the first stage and one in the second stage. Both stages are returned to Earth by parachutes.

The NK-33 engine is the highest performance LOx/kerosene engine available. It has been extensively tested and was fully qualified for the Russian lunar program. It incorporates an Aerojet-developed state-of-the-art electronic controller, ignition system, restart capability, an electromechanical actuator control valve, and a gimbaling system. The verification engine modifications are complete, and six tests have been completed. The NK-43 was developed for altitude performance and is essentially an NK-33 with an increased nozzle expansion ratio.

Kistler has teamed with Rocketplane Ltd., Inc., of Oklahoma and will continue operations as Rocketplane Kistler. The Rocketplane Kistler team expects to provide unique suborbital and orbital commercial space transportation services for passengers and cargo through its fleet of highly reliable, cost-effective, and reusable aerospace vehicles. Kistler’s successful restructuring should enable the company to complete the first K-1 vehicle, currently 75 percent complete, and conduct its first launch in 2007.

All license agreements are in place. There are 37 NK-33 and 9 NK-43 engines at Aerojet (equivalent to 180 missions) with ~40 engines still in Russia contracted to Aerojet. There is also a complete NK-33/43 engine design package at Aerojet with a technical support agreement in place and active. The Kistler K-1 vehicle, with its recoverable first and second stages and its available first- and second-stage engines, is an option as part of an advanced overall architecture for assured U.S. access to space.6

FALCON Small Launch Vehicles

As discussed earlier, a major element of a transformed total access-to-space architecture is the introduction of vehicles for ORS early in the far term of AFSPC’s SMP FY-06. Responsive spacelift is shown in the DoD space transportation roadmap in Figure 4-1. In the demonstration phase, 2006 through 2009, the objective is to have two or three small launch vehicles be flown. Some of the vehicles initiated under FALCON are expected to transition into cost-effective commercial launchers that could replace high-cost small vehicles.

Background

The DARPA/Air Force/NASA FALCON program started in August 2003. The overall goal of the program is to develop and validate in-flight technologies that will enable both near-term and far-term capabilities to execute time-critical, prompt global-reach missions while at the same time demonstrating affordable and responsive spacelift. The fundamental underpinning of the FALCON program is the belief

6

For additional information, see the official Kistler Web site at http://www.kistleraerospace.com. Last accessed on August 8, 2006.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

that a common set of technologies can be matured in an evolutionary manner that will provide a near-term (2007-2010) operational capability for responsive, affordable spacelift and prompt global reach from a small satellite (smallsat) launched from the continental United States (or equivalent reach from basing outside of the continental United States). These technologies might also enable future development of a reusable hypersonic cruise vehicle (HCV) in the far term (circa 2025).7,8

There are two tasks in this program. Task 1 is focused on a small launch vehicle (SLV) and Task 2 on a hypersonic technology vehicle (HTV). Together, the capabilities of placing small satellites or payloads into LEO and performing HTV missions in a responsive manner are an important step in the evolution of ORS capabilities for the Air Force (DARPA, 2004).9

The technologies for the launch vehicles needed to place a small payload into LEO or to place an HTV at its insertion point have enough in common that the design for both missions is encompassed within Task 1 (DARPA, 2004). For the first part of the mission, the top-level requirements for the operational system for the SLV are 1,000 lb payload (with the potential for an increase) to a 28.5o circular orbit, 100 nm altitude (baseline orbit for concept comparison) LEO (Weeks et al., 2005).10 The vehicle would have low recurring costs (less than $5 million), reach alert status within 24 hr, and then launch within 24 hr.

For the hypersonic systems, the objectives are unpowered, maneuverable, hypersonic glide.11 The vehicle would carry up to 1,000 lb of payload and have a minimum range of 3,000 nm. The reusable HCV would be an autonomous aircraft capable of taking off from a conventional military runway and striking targets 9,000 nm distant in less than 2 hr.

The first 6 months of the program (Phase I) were for concept development and identification of technologies. Phase II, which followed, was divided into three parts. Phase IIA covers authorization to proceed to preliminary design review (PDR). Phase IIB covers PDR to critical design review (CDR), and Phase IIC covers CDR to the demonstration flight of the small satellite spacelift mission.12 In Phase IIC, final flight hardware will be built up and flown no later than FY08. Demonstration flights are expected to carry prototype autonomous flight safety systems and low-cost tracking and data relay satellite system transceivers.13

After the FALCON program has been completed, DARPA will hand over the demonstration vehicle systems aspects to the AFSPC for development and implementation of the operational system. It is possible that the winning vehicles can contract directly with NASA or private entities (e.g., academia, amateurs, and other government agencies) to arrange for commercial launches.14

Four Vehicle Concepts in Demonstration Phase in 2005

Figure 4-5 summarizes the four vehicle concepts in the demonstration phase in 2005. Each concept is discussed in more detail in Appendix E.

7

For additional information, see http://www.darpa.mil/body/news/2003/falcon_ph_1.pdf. Last accessed on March 30, 2006.

8

David Weeks, NASA Marshall Space Flight Center (MSFC), personal communication to committee member Ivett Leyva on May 18, 2005.

9

Ibid.

10

Ibid.

11

For additional information, see http://www.darpa.mil/body/news/2003/falcon_ph_1.pdf. Last accessed on March 30, 2006.

12

David Weeks, personal communication to committee member Ivett Leyva on May 18, 2005.

13

Ibid.

14

Ibid.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

It is worth mentioning that SpaceX has its own funding. DARPA funds only the demonstration flight and making the launch operations responsive. The vehicle has been under development for a couple of years but was lost as a consequence of fire during its maiden launch in 2006.15

FALCON is the first concrete program devoted to the realization of affordable and responsive spacelift. Each of the four contractors for Phase IIA, Task 1, worked on several technologies to meet these goals. Some of the key technologies being developed or optimized that could be modified for or transferred to other programs are ablative thrust chambers, pressurization systems (VaPaK, Tridyne, etc.), low-cost avionics, hybrid combustion (using a patented staged-combustion concept), and composite tanks. Ablative thrust chambers were used by at least two contractors as an alternative to actively cooled ones. Composites were also being looked at by at least two contractors to reduce weight. Hybrid combustion was being revisited via a patented staged-combustion concept to achieve both combustion stability and performance comparable to that of liquid-fuel rockets. Common to all contractors is the objective of low-cost operations, which drove all of them to find innovative ways of streamlining their manufacturing, integration, transporting, and storage processes.

Another observation is that the subsystems of all the vehicles are modular and should be scalable, although to different degrees. These qualities potentially make some of the vehicle first stages candidates for replacement strap-ons for the Atlas or Delta vehicle families. Also, scaled versions of some of the first-stage engines could potentially be used as larger-thrust second-stage engines. FALCON promises to foster a new approach that designs space launch vehicles for versatility from the very beginning.

FIGURE 4-5 Vehicle characteristics for the four contractors in Phase IIA for DARPA FALCON program. SOURCE: DARPA (undated).

Programmatically, the FALCON approach has a lot of advantages. For example, the requirements were very few but very concise. This truly allowed for outside-the-box thinking, which is evident in the systems designed by the four contractors in Phase IIA, Task 1. The FALCON program also encourages

15

For additional information, see the official SpaceX website at http://www.spacex.com. Last accessed on November 7, 2006.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

small companies to enter the space access business. Each of the four companies under Task 1 has committed to at least two new technologies to reduce the cost of access to space. In all cases, the system configurations being developed must allow scalability, to mid- and even heavy-lift vehicles. They entail also, to different degrees, modular rack-and-stack approaches. The FALCON program provides opportunities to eliminate technology risks for larger launchers. In its early phases it stimulated the design and demonstration of new, low-cost, responsive technologies for space access.


Finding 4-5. The FALCON program is an initial response to the need for low-cost, operationally responsive access to space. This program plans to perform in-flight validations of technologies leading to highly responsive vehicles that can carry out time-critical, global-reach missions. The cost goal for FALCON-technology-based designs is $5 million (2003 dollars) per launch. Current costs for similar payloads using available small and medium-size vehicles are $20 million to $30 million. Successful FALCON demonstration vehicles and, later, production vehicles would open the door to a larger market for commercial space payloads. An increased launch rate would allow for the increased production of SLVs, which in turn would lower the cost of the vehicles through true mass manufacturing. Also, if more satellites could be launched each year, they would not need to be designed for a 5-10 year lifespan but could instead be updated or replaced more often. In FALCON, cost is prized over performance.

Expendable vehicles using low-parts-count, pressure-fed liquid propulsion systems such as systems used for the AirLaunch FALCON demonstrator and the SpaceX vehicle can be developed for much less money than reusable ones. Depending on the annual flight rate, they can also cost less per flight.


Recommendation 4-5. In September 2005, DARPA downselected to just one company for Phase 2B. DARPA should continue to fund and monitor this company to completion of the FALCON program objectives. The Air Force should evaluate the propulsion technologies to be demonstrated for the air-launched FALCON vehicle and include them in total system studies of options for ORS vehicles.

Air-Based Vertical-Launch Concept

As stated above, the overall goal of the FALCON program is to develop and validate, in flight, technologies that could provide both a near-term and a far-term capability to execute time-critical, prompt global-reach missions from the continental United States (or equivalent reach from basing outside the continental United States) while also demonstrating affordable and responsive spacelift for a variety of small satellites. Achieving these capabilities is an important early step in the evolution of ORS and global strike capabilities for the Air Force.

In the fall of 2005, another vehicle launch concept was disclosed that appears to have good potential for achieving some of the above capabilities—very fast, precision global and tactical strike and responsive cost-effective launch of satellites at the lower end of the smallsat spectrum16 into various LEOs (Smith, 2005). The concept grew out of efforts to find solutions to the severe time and geographic constraints associated with ground-based, boost-phase ballistic missile defense. The idea is to install a vertical launching system in a large-body aircraft. Such an aircraft could be on-station anywhere in the world it is needed. The system is illustrated in Figure 4-6.

16

Defined here as nano, micro, and <200 lb.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-6 Air-based vertical launch system (ABVL). SOURCE: Smith (2005).

The launcher subsystem is completely removable from the aircraft. Installation of flight-ready missiles into the self-sufficient module and installation of that module into the large-body aircraft would be carried out in separate ground facilities.

For satellite applications the separate compartments can be of various widths to accommodate various launch vehicles. One possible launch vehicle for a 50-lb-payload smallsat that could be accommodated in a 747 is illustrated in Figure 4-7.

FIGURE 4-7 Space launched nanosatellite system, BAE Systems Model SP-1a. SOURCE: Smith (2005).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

The feasibility of an ABVL is being studied by BAE Systems and ATK under a small DARPA contract. Beyond that, an integrated total systems engineering program would be necessary to establish propulsion requirements that can exploit the potential for responsiveness and low cost offered by such a system for small satellites. (The concept is also of interest for prompt-reach missile applications. See Chapter 5.) New technologies or modifications of existing stage or missile booster designs critical to meeting those requirements can then be identified, specified, and demonstrated.

Much of the launch dynamics and environment of an ABVL is very different from that of a ground-based launch, a Pegasus air launch, or the candidate FALCON vehicle Airlaunch, described above. Launching at high altitude (say, 45,000 ft) eliminates a large part of the Gt losses and a large part of the acceleration losses associated with ground launch (up to 3,500 ft/sec, depending on launch altitude and thrust/weight (T/W) ratio during the early part of the boost).17 The normally high L/Ds of launch vehicles are driven by the need for low-altitude drag and stability. Current large-body aircraft could accommodate many of the existing launchers for small missiles. However, this configuration may not fully exploit the launch concept given the individual space launch vehicle gross weights that could be supported in optimized launcher module volumes and the cargo capacity of the aircraft. On the other hand, the exit horizontal drag dynamics and subsequent pitch over to the optimum trajectory angle of attack would be heavily influenced by the launch vehicle’s diameter and by the specific thrust/time ratio and thrust vector control that could be achieved by the initial boost propulsion system.


Finding 4-6. Configurations for candidate launch vehicles (including parallel boosters or strap-on combinations), along with propulsion technologies such as propellant combinations (solids, storable liquids, gelled combinations, storable-oxidizer hybrids) and operating characteristics (including assured start-up profiles, thrust vector control, and rocket plume impingement patterns) need to be optimized to take full advantage of the potential new operationally responsive mission capabilities of aircraft-based vertical launch for small satellites, satellite arrays, and near-space military applications (see also Chapter 5).


Recommendation 4-6. The Air Force and DoD should sponsor a detailed system engineering study to fully understand the transformational potential of cost-effective, operationally responsive launch of small, micro-, and nanosatellites (particularly for large-number satellite arrays) utilizing air-based vertical launch concepts. The propulsion technologies that are needed to take full advantage of such launch platforms should be identified and developed.

Multimission Modular Vehicle Air-Based Launch

A new multipurpose airframe design is under consideration for potential future applications by DARPA. Such an airframe, currently designated the multimission modular vehicle (MMMV), would have a detachable centerline payload. It might have a joined-wing configuration to support the centerline that would enable this concept. The airframe is designed in such a way that the centerline payload could be either a passenger- or cargo-carrying fuselage that could also be equipped with folded rotor blades for emergency separation or self-transport.

The MMMV concept could provide a transformational access-to-space capability for future medium to large satellites. The aircraft could be configured to transport rocket-powered, access-to-space vehicles to high-altitude launch points at ideal geographic locations. As described above for smallsats, this would afford tremendous flexibilities in launch time, azimuth, or orbital inclination for large satellites. Also, a specialized missile pod could be used for tactical, strategic, or even antiballistic missile defense, enabling a more rapid response to emerging threats more rapid than is available today. The versatility of a

17

Losses of Gt are the result of the reduction by gravity of the vehicle acceleration produced by the vertical component of the engine thrust vector integrated over the angle of attack during the launch trajectory.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

removable centerline payload is the key to this operational flexibility. The basic lifting aircraft configuration is illustrated in Figure 4-8. An MMMV carrying a medium to large spacecraft launch vehicle is depicted in Figure 4-9.

FIGURE 4-8 MMMV concept. SOURCE: NASA MSpFC.

FIGURE 4-9 MMMV concept for medium to large spacecraft. SOURCE: NASA MSFC.

Finding 4-7. As discussed above for air-based vertical launch of small launch vehicles and missiles, candidate medium to large launch vehicle configurations and their propulsion technologies would need to be optimized to take full advantage of the potential for a modular configuration aircraft to transform mission capabilities by enabling high-altitude launch. Some of the propulsion technology aspects that need to be investigated include propellant combinations capable of long on-station standby (solids, storable fuels and oxidizers, gelled combinations, hybrids); first-stage chamber pressures and expansion ratios; and various operating characteristics, including assured start-up profiles, thrust to weight profiles, thrust vector control, and rocket plume impingement patterns.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

Recommendation 4-7. The Air Force and DoD should combine a detailed system engineering study of the multimission modular vehicle air-based launch system for medium-sized vehicles with the study of air-based vertical launch for small vehicles called for in Recommendation 4-6. Air Force and DoD sponsorship would ensure they are focused on Air Force and DoD criteria for optimizing mission success. The study would identify propulsion technologies (modifications or new concepts) that should be evolved in order to take full advantage of such air-based launch platforms for both strategic and operationally responsive missions.

Operationally Responsive Spacelift Requirements

The DoD’s Space Science and Technology Strategy states that assured access to space is the highest priority within the space support mission area, and a responsive space capability is directly coupled to both the space support and force enhancement mission areas (DoD, 2004). An important element of a transformed total access-to-space architecture is the introduction of ORS vehicles early in the medium term of the SMP FY06. Some of the missions driving the ORS architecture are indicated in Figure 4-10. In the Air Force’s roadmap of ORS spirals (Figure 4-1), selected vehicles from the FALCON program, described above, would continue developmental and operational flights as part of the Air Force’s fleet of small launch vehicles into the far term (James, 2005). Each of the selected concepts would probably evolve into a family of fast-response expendable vehicles that could launch payloads from 2,000 to 10,000 lb to LEO. The roadmap also shows the start of full-scale development of an ORS vehicle in 2010.

FIGURE 4-10 Missions in operationally responsive spacelift. SOURCE: Hampsten et al. (2005).

The objectives for ORS vehicles are shown in Figure 4-11. Meeting these objectives may necessitate a number of new propulsion subsystem technologies in addition to applicable existing qualified subsystems.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-11 AFSPC operational objectives. SOURCE: Hampsten et al. (2005).

As discussed in some detail in the first section of this chapter, an integrated total systems engineering process in which propulsion requirements for these vehicles are established and technologies critical to meeting those requirements are defined is crucial to the success of any new launch-to-space or in-space vehicle program. Using “mission success” as the primary selection criterion for this systems engineering process provides a powerful quantitative tool for the design of low-risk, cost-effective ORS concepts for Air Force future needs.

For a specific mission defined by a set of requirements issued by a user program authority, “mission success” 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 program. In a total systems engineering process, “mission success over the required life of the system” can be the primary criterion for selection from among total configuration options. “Total configuration” encompasses a system’s overall architecture, including all of its major flight system elements, all of its required direct supporting elements, its required logistics architecture, and the supplier base.

For example, the configuration of the first stage of a new operationally responsive launch vehicle, including the design of its propulsion elements, would not be selected using conventional (but subjective and ineffective) criteria such as lowest weight. Instead, it would be selected by virtue of being an element of a particular total configuration option. “Mission success” would be characterized for each total system option as: “achieving the functional capability wanted with a specific schedule uncertainty profile (using a consistent methodology), a specific total-life-cycle price uncertainty profile (using a consistent methodology), and within the program authority accepted (not the lowest) quantitative risk uncertainty level profiles for functional performance and total systems operational reliability.” The options could then be selected based on the most advantageous mission success profile.

Some of the most critical design options for first-stage propulsion (e.g., expendable vs. reusable; storable vs. cryogenic propellants; thrust/ time ratios; and pumped vs. pressure fed) are completely entwined with the total system architecture, including such things as fixed, distributed, or mobile launch facilities, active mission launch rate, and yearly launch rates. Such characteristics would benefit from not being locked in before an objective total systems engineering process has been completed, including the validation status of the design criteria for all proposed critical technologies.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

This is the dominant factor in making objective evaluations of the schedule and cost risks and for development engineering of the operational and life-cycle cost risks of the propulsion systems. It also permits quantitative and consistent comparisons of concepts across the broad trade-off space of propulsion systems. Propellant selection, for example, would require trades between pumped vs. pressure fed; pressurization subsystems for net positive suction head or propellant feed; optimization of chamber pressure and nozzle expansion ratios for first or second stages; expendable vs. reusable first and/or second stages; metals vs. composites for tanks or motor cases; ablative vs. cooled combustion chambers.

Most important, when rigorously applied, such a program would eliminate the identification of unvalidated design criteria associated with technologies for critical elements of conceptual propulsion systems or for upgrades of existing subsystems proposed for various candidate vehicles. These unvalidated system element design criteria (which include criteria for the element’s total operating environment) are the primary drivers of a development program’s engineering, operational, and cost risks. (In a number of past cases, the use of designs for which too few criteria had been validated was the first cause of catastrophic failures.)

Affordable Responsive Spacelift Vehicle

The Air Force has set up a program to demonstrate a subscale vehicle and validate the system concept. Called affordable responsive spacelift (ARES), the vehicle will be evolved into the ORS family of vehicles.

The ARES program evolution spiral was planned to start in 2005. The Air Force has been working on conceptual systems engineering for ARES and has completed an initial group of analyses. The result of the Air Force’s current analysis is a basic architecture concept for a reusable, fly-back-to-launch-site, rocket-engine-powered first stage and an expendable, rocket-engine-powered second stage.

The Air Force believes the ARES hybrid is the medium-term solution for a revolutionary spacelift capability (James, 2005). It also foresees that an ARES flight demonstration in 2010 will provide confidence in full-scale system costs and operability and will enable fielding a system in an affordable fashion.

In the committee’s opinion, if the ARES system design is to be selected via a total systems engineering process to provide confidence in full-scale vehicle development, it must essentially lock in most of the critical technologies for the full-scale configurations by default. To proceed confidently with competitive conceptual designs for a subscale demonstrator starting in 2005 and implementation of a selected configuration development program starting in 2006, the choice of propulsion technologies would benefit from being constrained by a total systems engineering process whose elements have been qualified or at least have extensive validating data.

The committee’s review did not turn up any transformational or revolutionary technologies mature enough to be considered for ARES (and therefore for any full-scale vehicle that would be “justified” by the subscale demonstrator). Also, the committee could identify only two existing rocket engines that might meet the propulsion system requirements for the ARES hybrid: the Aerojet AJ26-58/59, which would constrain the first stage to LOx/RP-1, and the RL-10 family, which would constrain the second stage to LOx/LH2.

New rocket engine designs such as those described below would have to be objectively compared to the Aerojet AJ26-58/59 and the RL-10 family. Only one of the engines has significant test data available. Total systems engineering dictates that all engine comparisons be based on rigorous objective assessment of forecasted performance, development cost and risks, operational characteristics, and logistic support required for new engine designs.

There could be many opportunities to be clever in configuring total access-to-space systems architectures for fixed and mobile (including air-based) launch locations, payload module integration, launch operations associated with both reusable and /or expendable vehicles. industrial support, parts and propellant storage and logistics, and so on. However, as stated above, the committee does not see any revolutionary technologies that could be incorporated into the vehicles’ propulsion systems. Tanks and

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

feed systems may incorporate design improvements and eliminate known failure modes or improve margins, but they do not constitute revolutionary or transformational technologies. The propellant combinations that are realistically available to ARES are well known and have been in use for a long time. Based on “mission success” total system design criteria, the ORS missions out into the early part of the far-term (FY18-FY30) do not require revolutionary technologies to accomplish. In fact, at this point in time, the various risks of committing to unvalidated technologies are much greater than any overall gain claimed for system performance.

The Air Force recognized this real situation and stated as follows (Hampsten et al., 2005, p. 12):

Remember … ARES-SD is the first phase of an acquisition program. [The] Air Force wants to achieve its goals using the lowest risk approach practical The ARES management team uses the term technologies in the generic sense of describing the technological means to an end. It is not intended to indicate specifically immature, “high-tech,” stretch design goals, or high-risk technologies. This is not a tech-push effort.

The committee concludes that ORS missions out into the early part of the far-term (FY18-FY30) will not have (and in the committee’s opinion do not need to have) revolutionary propulsion technologies. In fact, the various risks of committing to unvalidated technologies at this point in time are much greater than any potential gain in rocket propulsion system performance. If there is to be a revolutionary ORS capability in the medium term, it will be created by very innovative total systems architecture and operations processes and by high margins against retained failure modes, not by revolutionary rocket propulsion system technologies.

INITIATIVES TO ESTABLISH NEW PROPULSION TECHNOLOGY BASE

National Aerospace Initiative

The National Aerospace Initiative (NAI) began in 2001 as a joint technology initiative by the DoD and NASA. It is not, however, a program for system development or acquisition (NRC, 2004). In fact, “the goals of NAI are to renew American aerospace leadership; push the space frontier with breakthrough aerospace technologies; revitalize the U.S. aerospace industry; stimulate science and engineering education; and enhance U.S. security, economy, and quality of life” (NRC, 2004, p. 3). The initiative rests on three pillars: high-speed, hypersonic flight, access to space, and space technology. An NAI executive office was created to foster collaboration between NASA and DoD and develop goals, plans, and roadmaps for the three pillars. The idea was for NAI to start by identifying the capability objectives for the future systems; to use the goals, objectives, technical, challenges, and approaches (GOTChA) process to analyze the technology development challenges and options; to establish investment plans; and, finally, to coordinate the combined efforts of the involved parties to execute the developed technology plans (NRC, 2004).18 The elements of the NAI are shown in Figure 4-12.

18

“The GOTChA process combines layered analysis (goals are analyzed to determine objectives, which are analyzed to determine technical challenges, and so on) and planning (projects are identified and roadmaps developed to address the challenges)” (NRC, 2004, p. 11).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-12 Technology framework for National Aerospace Initiative. SOURCE: Sega (2005).

During 2002, the NAI goals, definitions, and plans were further developed. In 2003, at the request of the Air Force and DDR&E, NRC empanelled the Committee on Review and Evaluation of the National Aerospace Initiative, which produced the report referred to here (NRC, 2004). That committee looked at only the first two pillars of the NAI, high-speed, hypersonic flight and access to space, and was tasked, in part, with the following:

To assist the Department of Defense (DoD), the services and agencies, and NASA by providing an independent evaluation of the feasibility of achieving the science and technical goals as outlined in the National Aerospace Initiative, the National Academies, under the leadership of the Air Force Science and Technology Board, will form a committee to answer the following general questions concerning the NAI:

  1. Is NAI technically feasible in the time frame laid out?

  2. Is it financially feasible in the same time frame?

  3. Is it operationally relevant? (NRC, 2004, p. viii)

The NRC’s NAI review committee agreed with the general goal of demonstrating technologies that greatly increase space access and reliability while reducing cost. However, the committee did not believe that all the payoffs would be realized in the announced time frames. It also found the access to space pillar underfunded. Specifically, NAI had envisioned a multiphase demonstration program (2008 and 2015) with increasingly capable reusable rockets, which in the committee’s view was underfunded. The review committee found NAI operationally relevant, especially with respect to capability goals and missions such as ORS.

The review committee also looked at the activities of IHPRPT when information was available to learn whether there were synergies and commonalities with the goals of NAI. However, NAI and IHPRPT are separate programs and do not interact directly. According to the committee, the critical near-term technologies for the space access pillar are these:

… advanced materials for use in propulsion and thermal protection systems; integrated structures; electrical/hydraulic power generation and management technologies; software transportability; and error-free software generation and verification. Furthermore the development of computational analysis tools and methodologies should be emphasized— especially when coupled to test analysis and ground test facilities (NRC, 2004, p. 6).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

It recommended that in the area of propulsion, research be done on engine reusability and reliability, which should include high-strength, LOx-compatible materials, and also materials that withstand hydrogen and hydrocarbon combustion. Other recommended material research areas were durable lightweight thermal protection, structural materials, and reusable propellant tanks (NRC, 2004).

The NRC’s NAI committee suggested that NAI integrate some of the available advances in intelligent sensors and thermal control components. Another recommendation was to focus on lowering the cost of aerospace software production. Attention should also be paid to vehicle health management technologies that increase safety and engine life and decrease maintenance costs (NRC, 2004).

According to the review committee, more aircraftlike operations require development of more robust vehicles, more efficient ground operations, and automated flight planning. If necessary, payload capability could be sacrificed for design robustness (NRC, 2004).

The goals and direction of NAI changed on January 14, 2004, when President Bush announced a plan to develop and test a new crew exploration vehicle by 2004, human missions to the moon (circa 2014), and, later, missions to Mars (NRC, 2004). This announcement was made after the committee had submitted its draft report for external peer review. Some of the changes that have occurred since 2004 can be seen in the evolution of the NAI roadmaps (Figures 4-13 and 4-14). As can be seen, the access to space and space technology pillars of NAI continue to evolve.

NAI represents cooperation, better utilization of resources, and maximization of synergies. However, it is hard to identify how the money from the NAI initiative is being spent beyond the first layer of general funding. The recommendations and observations of the NRC’s NAI committee, summarized above, still appear to be valid for the present study.

FIGURE 4-13 Updated NAI roadmap. SOURCE: Sega (2005).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-14 Test/demonstrator roadmap—high speed/hypersonics and space access, 2004. SOURCE: NRC (2004).

Integrated High-Performance Rocket Propulsion Technology

The IHPRPT program was initiated in 1994 and has been in place for 12 years. It is a joint effort of government and industry focused on affordable revolutionary technologies for reusable, rapid-response, military global-reach capability. It addresses sustainable strategic missiles, the trade-off between long life and increased maneuverability, spacecraft capability, launch vehicle propulsion, and high-performance tactical missile capability. IHPRPT attempts to emulate the metrics of the integrated high performance turbine engine technology (IHPTET) program, which succeeded in developing and testing new turbine engine technologies.

Although funding for IHPTET has been severely limited, contractors have had considerable freedom to develop new technologies that can improve the performance and life of both solid and liquid rocket engines. Several contractors say the main difficulty is that there is no clear definition of Air Force needs.

Specific performance goals have been established for each element of the program, some of which are summarized in Table 4-2.

TABLE 4-2 IHPRPT Goals for Improvements in Boost and Orbit Transfer Propulsion

Goal

2000

2005

2010

Reduce stage failure rate

25%

50%

75%

Improve mass fraction (for solids)

15%

25%

35%

Improve Isp (sec)

14

21

26

Reduce hardware costs

15%

25%

35%

Reduce support costs

15%

25%

35%

Improve thrust to weight (for liquids)

30%

60%

100%

Mean time between removal (mission life-reusable in number of missions)

20

60

100

SOURCE: Huggins (2005).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

The building block approach to planning for technology insertion seems to be working well, and there appears to be excellent cooperation between the Air Force and the contractors and between the Air Force and NASA. IHPRPT provides a means for government and industry to evolve along a common path.

Another significant limitation is that component testing and validation are not sufficiently funded by the government. The uncertain outlook for commercial launch opportunities produced by this underfunding, together with the aforementioned failure of DoD to specify needs, has discouraged the large rocket propulsion companies from investing their own decreasing resources in new propulsion subsystems.

Air Force Research Laboratory Efforts Under IHPRPT

Combustion Stability

The Air Force Research Laboratory (AFRL) is planning to upgrade its combustion stability models by including more physics in their development as well as their validation. The effort is not currently funded, although funding is being actively sought.

Nonequilibrium Flow

A number of programs are being conducted to validate direct simulation Monte Carlo models, examine advanced micropropulsion concepts and plume-spacecraft interactions, and model nonequilibrium flows in very small nozzles. A major effort is being undertaken to model capillary discharges.

Combustion Devices

A number of programs are being carried out in this area, including measuring the patterns of injectors, examining the mixing of supercritical flows, and developing injector design methodology. In addition, a high-heat-flux facility has been developed to measure the decomposition of some new and promising hydrocarbon fuels.

Solid Propellants

In-house testing of insulation materials and new oxidizers for solid propellants has been concluded, and current efforts are focused on supporting the land-based strategic deterrent. To this end, facilities to formulate and test new propellants being developed by industry are being upgraded.

Liquid Engine Technology

New hydrocarbon propellants that have been synthesized by laboratory chemists are to be tested in a small engine of about 1,000 lb thrust. The intent is to measure a variety of parameters. A basic effort is being undertaken to determine the effect of channel aspect ratio on the ability to cool rocket chambers. The effect of curvature on cooling ability will also be measured.

Materials Applications

Nanophase aluminum has better properties than standard aluminum. However, the consolidation of nanophase aluminum particles usually results in the growth of the grain boundaries and the loss of the improved properties. AFRL is attempting to consolidate these particles without inducing grain growth.

The relationship between structure and properties when polyhedral oligomeric silsesquioxane (POSS) is included in polymers is being investigated. POSS is being examined to see if it can be used as a

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

coating for solar cells that would allow them to survive in space. The intent is to see if the POSS would form a layer whose cracks would be self healing under the influence of atomic oxygen.

AFRL has developed a technique to densify carbon and carbon materials using high-carbon-content liquid materials along with a catalyst. This reduces the densification process time by a factor of 4 or more. Studies are ongoing to further understand how this technique carbonizes to improve the process.

Organometallics have been used in supercritical carbon dioxide (CO2) to produce superior metal coatings on various substrates. A current study is attempting to determine the process parameters to characterize and improve the coating process.

Propellant Development

A significant amount of work has been conducted by AFRL at Edwards Air Force Base on the preparation of green propellants to replace hydrazine. These mono- and bipropellant materials promise to provide high-density impulse, surpassing the performance of the most commonly used toxic monopropellants and in some cases approaching that of toxic bipropellants. Current work assumes the use of hydrogen peroxide as the oxidizer. Essentially no work has been done on new oxidizers.


Finding 4-8. The AFRL Space and Missile Systems Division is undertaking a variety of interesting and potentially valuable in-house programs. It appears to be developing technology that will be very useful, such as predicting the existence of certain energetic compounds and their synthesis, determining the coking properties of hydrocarbon propellants, and developing combustion instability models. Unfortunately, there does not appear to be much in-house work in the liquid engines and solid motors areas. The more basic work seems to be of high quality, but its basic nature and not knowing where the Air Force wants to be in the future make it very difficult to set the priorities for these efforts or even determine if they are the best ones to undertake. A thorough review by outside experts might help in prioritizing the efforts.


Recommendation 4-8. The Air Force should develop in-house test beds for liquid, solid, and hybrid rocket motors. Because limited funding seems to be at least part of the reason this is not being done, the Air Force should seek to increase the funding for both liquid and solid rocket test beds at AFRL.

Contractor Efforts Under IHPRPT Funding

Pratt & Whitney

Pratt & Whitney was one of the first companies to become involved in IHPRPT, gaining three programs early on. One program was to develop a liquid hydrogen turbopump, the second was to develop an expander cycle combustion chamber, and the third was to combine these two into an advanced upper stage demonstration. Other components were being provided from other sources.

The liquid hydrogen turbopump was developed and tested more than 19 times, including several times when rotation was achieved. Unfortunately, the turbopump was damaged during some of the testing, as was the combustion chamber that had been developed. Because the program was hardware poor, the demonstration never came about. The two programs did, however, provide significant technical information that was fed into other IHPRPT programs.

Rocketdyne

Rocketdyne has focused its IHPRPT efforts on improving the performance and reliability of large liquid rocket propulsion systems, with emphasis on turbopumps and combustor reliability.

In the turbopump area, the company has had an effort on hydrostatic bearings for many years and now feels confident that it has the design tools to develop reliable turbopump bearings over a wide range

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

of sizes. Hydrostatic bearings have several advantages over conventional rolling-element bearings, including degreased coolant flow and longer life. Hydrostatic bearings were recently tested for a few seconds in the integrated powerhead demonstrator (IPD) (jointly with NASA), but much more testing is needed before they can be certified for flight. Rocketdyne has also pursued several other turbopump-related efforts, including turbine blade damping to reduce the risk of fatigue damage to blades and advanced high-strength materials for rotating machinery. The latter, if successful, would allow significantly higher speeds, particularly in smaller turbopumps.

There has also been work done on gas and gas main injector technology and oxygen-rich preburner technology, both of which offer performance improvements but require the development of oxygen-compatible materials. In the latter arena, high-specific-strength materials that are also oxygen compatible are required, and this has been a major challenge. However, significant progress has been made, and strength improvements of 40 to 50 percent have been demonstrated.

Development of improved design tools has been a major focus at Rocketdyne, with the emphasis on computational fluid dynamics (CFD) and thermal modeling. Some of these improved design tools were used in the development of the RS-68 (Delta IV) engine, and Rocketdyne estimated that the development cost of the RS-68 was reduced by a factor of 3 through the use of these tools. It further estimated that a sixfold cost reduction could be attained if critical skills and experienced staff could be retained for the next cycle of engine development programs.

Rocketdyne said that a big problem was that it had no clear picture of future Air Force needs in rocket propulsion. Rocketdyne management set a high priority on the retention of expert knowledge and on a ground demonstrator engine for testing advanced components.

Aerojet

Aerojet divides its IHPRPT efforts into three areas: liquid propulsion, solid propulsion, and in-space propulsion. Each comprises a number of programs; the text below focuses on liquid propulsion and solid propulsion.

Liquid Propulsion. The programs in this area are for upper-stage engine technology (USET), liquid crystal polymers, the advanced lightweight chamber and nozzle, the reaction control engine (RCE), and the IPD:

  • Upper-stage engine technology. This program, which has been under way since November 2003, aims to develop physics-based design tools and methodologies to replace empirically based design tools and to develop LH2 turbopump assembly hardware for tool validation that supports IHPRPT goals and technology insertion to replace the EELV RL-10. The program will shortly complete the CDR stage.

  • Liquid crystal polymer. During the nearly 3 years that this program has been under way, it has been attempting to develop a liquid crystal polymer/composite braiding material system for a lightweight, low-cost duct that will meet the demanding environments in next-generation cryogenic engines.

  • Advanced lightweight chamber and nozzle. This program aims to design, fabricate, and test hardware that can be used for four purposes: (1) to demonstrate single-bell and dual-bell operation in 40 K hot-fire testing at sea level, throttling the engine to determine flow field characteristics, and to anchor a CFD model; (2) to develop and demonstrate a high-temperature hot gas wall coating, to satisfy the IHPRPT goal of 60 missions, for an RP-1 regeneratively cooled thrust chamber; (3) to demonstrate the effectiveness of multipoint film cooling and implement such cooling; and (4) to build a subscale titanium nozzle to demonstrate key manufacturing processes and design features.

  • Reaction control engine. This relatively new program has the objective of demonstrating a TRL of 6 for LOx/ethanol auxiliary propulsion system.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×
  • Integrated powerhead demonstrator. The objective of this program is to provide the combustion devices for the Air Force’s IPD 250,000 lbf, LOx/hydrogen, full-flow staged combustion engine and test them at Stennis Space Flight Center. Several successful second ignition tests have been performed.

  • Thrust augmented nozzle. This program aims to fabricate and test thrust augmented nozzle hardware to demonstrate the percent thrust augmentation and system benefits, including engine system thrust/weight and thrust/volume, engine operating pressure, and altitude performance.

Solid Propulsion. The programs in this area are Atlas V; IHPRPT Phase IIB Demo (including SRM modeling and energetic propellants); sensor application and modeling; advanced second stage; FALCON small launch vehicle, and rocket system launch program flexseal.

  • Atlas V. The objective of this program is to design, develop, and produce low-cost solid rocket boosters (SRBs) to support the EELV Atlas V launch vehicle program. The development of these motors was qualified in 2003. There have been three successful flights to date. Between 7 and 15 SRBs are expected to be produced every year through 2011. Currently a Block B upgrade qualification is under way. It addresses material obsolescence and incorporates robust nozzle features.

  • IHPRPT Phase IIB demonstration motor. To prepare for future booster motors expected to play a primary role in military, commercial, and NASA missions, this program is updating a small ICBM second stage with a new Class 1.3 HTPB propellant and composite case and nozzle. It will provide an updated motor incorporating the best technologies to support future missile defense, space launch, and nuclear deterrent booster systems. The motor will offer improved Isp, longer life, higher T/W, and lower cost. It has a desubmerged lightweight nozzle, fewer parts and interfaces, a carbon-carbon exit cone, a wet-wound graphite/resin system, no dome reinforcements, strip-wound Kevlar ethylene propylene diene monomer (EPDM) insulation, a Class 1.3, 90 percent solids HTPB/RDX propellant, a consumable igniter, and a smaller electromagnetic thrust vector actuator.

  • Sensor application and modeling. This program seeks to avoid missile failures by using sensors to measure aging properties without adversely impacting the structural or chemical integrity of a motor. This will be accomplished by conducting a literature search and evaluation of all available sensors, followed by downselection to mature sensors. Then, inert-propellant, 5-in., instrumented composite motor tubes for laboratory-scale work will be built and cast with defects. Based on the results of this, an inert-propellant, 10-in., instrumented motor case will be built and cast with and without defects. Testing will then determine whether the sensors can identify the defects.

  • Advanced second stage. The program objective is to evaluate, develop, and demonstrate innovative solid rocket motor technologies in an advanced upper-stage configuration applicable to future advanced and affordable strategic strike systems. To accomplish this, the following steps will be taken: (1) conduct design trades/payoff analysis to evaluate technologies for a future second-stage system balancing cost and performance; (2) develop key solid rocket motor technologies and advanced manufacturing processes that are optimized on life-cycle costs to ensure an affordable future ICBM system; and (3) integrate technologies and processes in a full-scale demonstrator motor that will be tested at altitude operating conditions at the Arnold Engineering Development Center (AEDC). The program suffers, however, from severe underfunding in the areas of component testing and validation.

  • FALCON small launch vehicle. The objective of this program is to design a reverse-dome/forced-deflection nozzle and a gas injection thrust vector control (GITVC) system for the second stage of the FALCON small launch vehicle. This will be accomplished by designing the second-stage composite case, producing two reverse-dome/force-deflection

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

nozzles and two GITVC systems consisting of eight valves and one controller, producing two second-stage composite motor cases, and providing support for systems engineering and for ground and flight test vehicle engineering.

  • Rocket system launch program flexseal. This program converts surplus Minuteman II second-stage motors into first-stage suborbital launch boosters by modifying the fixed nozzle system so that it uses a flexseal movable nozzle and hydraulic actuation system, and it also provides technical support to ensure the thrust vector system meets evolving requirements for launch missions.

Northrop Grumman

Northrop Grumman is carrying out research and development across a broad spectrum of propulsion technologies. An important project related to access to space is the design of an advanced USET. Northrop Grumman was awarded a contract in September 2003 to develop tools and a full-scale modeling capability for upper-stage engines.

USET is a 5-year old IHPRPT program, funded and managed by AFRL at Edwards Air Force Base, and slated to end in FY08. It has the primary goal of improving the software design and analysis tools used for advanced rocket engine development. An informal objective is to improve the interconnectivity, efficiency, and optimization routines of a group of presently diverse softwares such that the engine design, analysis, and optimization process will take one-tenth as long as it now takes and require one-tenth as much labor. The basic concept is to use the best of currently available commercial or university software programs—e.g., CFD++, ANSYS, SINDA-Fluent, ROCKETS, TDK, Visual DOC, and Concepts NRC Agile Engineering package for turbopump assembly (TPA) analyses—to institute automated input/output exchanges between the various software packages (via a commercially available wrapper program), and to institute a design-of-experiments/automated optimization/automated sensitivity analysis capability (via a commercially available optimizer program). A very limited amount of customized software code is expected to be required, but some existing codes are being improved and updated under USET (e.g., pump cavitation modeling).

While the initial USET development focuses on a LOx/LH2 engine at 40,000 lbf, the contract specifically requires that the software tools be applicable to highly off-nominal run conditions, other propellants, different thrust levels, and other engine power cycles. At the end, this improved, coupled suite of design and analysis tools will be made available—under U.S. government control—to U.S. industry, academia, and DoD organizations to improve our nation’s competitiveness in developing new rocket engines. The schematic in Figure 4-15 represents a standard split expander cycle rocket engine using series-driven turbopumps. Engine components in this figure are the LOx turbopump, LH2 turbopump, regeneratively-cooled thrust chamber, split control valve, oxygen main valve, fuel main valve, and turbine bypass valves.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-15 USET upper-stage schematic. SOURCE: Northrop Grumman Corporation.

A new high-performance 40,000-lbf-rated LH2 turbopump will be built and tested to validate the improved suite of design and analysis software. Originally, an advanced regeneratively-cooled thrust chamber assembly was to have been built for a similar purpose, but this was struck from the program due to funding limitations. The only detail design work on the program is directed to the LH2 TPA, and the only testing on the program will be on the LH2 TPA. The budget allocated to Northrop Grumman’s space technology effort for completing the USET program was about $30 million. Additional information on USET can be obtained from AFRL at Edwards Air Force Base.

USET is an important step toward achieving the ultimate vision—computerized development of high-fidelity virtual engine designs that can be tested on a virtual test stand. It is anticipated that high-accuracy comparison of predicted performance with delivered performance, especially under transient and highly off-nominal run conditions, will become a reality. The funding and schedule constraints of the current USET contract preclude achieving this ultimate goal.


Finding 4-9. Funding the continued evolution of computerized high-fidelity designs for engines and propulsion systems that can be tested on a virtual test stand and then flown virtually could be one of the most cost-effective and highly leveraged investments that DoD and the Air Force could make. The potential for reducing the huge costs of cut-and-dried development of rocket engines and their associated propulsion systems is enormous when the time line is extended into the indefinite future. Compared to those savings, the expenses of making virtual rocket propulsion system design engineering a reality are almost trivial!


Recommendation 4-9. DoD and the Air Force should fund the continuing evolution and process validation of computerized high-fidelity virtual engine and propulsion system designs.

Other Efforts Under Government or Industry Funding: New Engine Designs and New Propellants, Feed Systems, Pressurization, and Materials

Four competing booster engine design and development programs were initiated and funded under the SLI at MSFC in 2001. Boeing Rocketdyne Power and Propulsion had two of the designs, the RS-83 (660,000 lb thrust LOx/LH2) and the RS-84 (1,050,000 lb thrust LOx/RP-1). TRW Space and Electronics

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

of Redondo Beach, California (now Northrop Grumman) offered a LOx/RP-1-fueled engine in the 1-million-pound-thrust class, which it named the TR-107. The Co-optimized Booster for Reusable Applications (COBRA) was a LOx/LH2-fueled engine in the 600,000-lb-thrust class that was to be designed and developed by a joint venture between Pratt & Whitney of West Palm Beach, Florida, and Aerojet of Sacramento, California.

RS-83 Engine, Boeing Rocketdyne

The Boeing Rocketdyne Propulsion and Power Company designed the RS-83 as a staged combustion LOx/LH2 main booster engine system with a simple fuel-rich preburner in place of the dual individual preburners on the SSME. Taking lessons learned from the first-generation reusable launch engine (the SSME is also manufactured by the Boeing Rocketdyne Propulsion and Power Company), the RS-83 engine was intended to be simpler to build and maintain and to be more controllable and reliable. Advanced design features included turbopumps with easy access and fabrication techniques using selectively net-shaped components made by powder metallurgy.

A conceptual diagram of the RS-83 is shown in Figure 4-16; the engine design performance and operating characteristics are summarized in Table 4-3.

FIGURE 4-16 Conceptual design of the RS-83 engine. SOURCE: NASA.19

19

NASA Stennis Space Center Propulsion Testing, https://rockettest.ssc.nasa.gov/ssc_ptd/images_sscptd/ssc_projects/rs83/rs-83_proe.jpg. Last accessed on September 19, 2006.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

TABLE 4-3 RS-83 Engine Key Design Characteristics

Characteristic

RS-83

SSME (Block II)

Thrust

 

 

sea level (lbf)

66,800

393,800 at 104%

vacuum (lbf)

749,600

448,800 at 104%

Isp, vacuum (sec)

445.7

452

Chamber pressure (psia)

2,800

2,994

Mixture ratio (sea level, altitude)

6.9:6.0

6.03:1

Engine T/W, sea level

55

73.12

Area ratio

40:1

69:1

Mean time between removal

(mission life-reusable in number of missions)

50-100 missions

100 missions

Catastrophic reliability

0.999958

0.9999

Throttling (% of thrust)

50-100

67-104 (can go to 109 for abort contingencies)

The program plan focused on the early development of critical engine components with the overall goal of identifying and reducing the risk associated with the development and testing of these elements. The engine design team identified five critical tasks to reduce component risk: (1) hydrogen-compatible materials, (2) turbine damping, (3) subscale liquid preburner, (4) electromechanical actuator (EMA) sector ball valve, and (5) integrated vehicle health monitoring safety and prognostic algorithms.

Unfortunately, NASA cancelled the RS-83 design and development program, along with the other three large reusable booster engine programs—XRS-2200, RD-146, and Fastrac and MC-1, mentioned above—that had been sponsored under the SLI when it redirected its program to reflect President Bush’s space initiative, amended in January 2004.

RS-84 Engine, Boeing Rocketdyne

The RS-84 engine program was proposed as the first U.S. reusable hydrocarbon-fueled, oxygen-rich, staged-combustion liquid rocket engine. One of the primary goals of the engine development effort was to develop a highly reliable and low maintenance cost engine as a part of NASA’s SLI for the next-generation reusable launch system. The Rocketdyne Propulsion and Power Division of the Boeing Company was awarded the contract to design the RS-84 prototype engine for NASA’s Next Generation Launch Technology (NGLT) program. The kerosene-fueled RS-84 engine was one of several technologies competing to power NASA’s next generation of launch vehicles.

The RS-84 engine development program was one of two competing efforts to develop an alternative to existing hydrogen-fueled engine technologies (e.g., SSME). The engine was to be fueled by kerosene, a relatively low-maintenance fuel with high performance and high density, meaning it takes a smaller fuel-tank to achieve greater propulsive force than other technologies. That benefit translates to more compact engine systems, easier fuel handling and loading on the ground, and shorter turnaround time between launches. All these gains, in turn, reduce the overall cost of launch operations, making routine space flight cheaper and more attractive to commercial enterprises. In addition, because it is not a cryogenic (extremely cold) fuel, like hydrogen, propulsion-related ducts, valves, lines, and actuators do not require insulation, saving weight and cost. Table 4-4 shows the proposed attributes of the RS-84 engine.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

TABLE 4-4 RS-84 Engine Key Design Characteristics

Characteristic

RS-84

SSME (Block II)

Propellants

LOx/RP-1a

LOx/LH2

Thrust

 

 

sea level (lbf)

1,064,000

393,800 at 104%

vacuum (lbf)

1,130,000

448,800 at 104%

Isp vacuum (sec)

324

452

Chamber pressure (psia)

2,800

2,994

Mixture ratio

2.7

6.03:1

Area ratio

20:1

69:1

Life

100 missions

100 missions

Throttling (% of thrust)

65-100

67-104 (can go to 109 for abort contingencies)

aRP-1, rocket propellant 1, a special grade of kerosene suitable for rocket engines.

COBRA Engine, Pratt &Whitney and Aerojet

The COBRA program was initiated to develop one of the engines being considered by the SLI for use on a next-generation reusable launch vehicle. The goal of COBRA was to produce a rocket engine prototype that would be simple to operate, provide high reliability and a long life, and reduce cost per launch by virtue of being reusable. COBRA planned to incorporate a reusable, hydrogen-fueled liquid booster engine with a thrust of 600,000 lbf. The engine was to be developed by a joint venture between Aerojet and Pratt & Whitney Space Propulsion.

The COBRA design consisted of a single fuel-rich preburner, staged-combustion engine using LOx and LH2 as propellants. The engine was to be designed to have a 100-mission life span with a 50-mission maintenance check-up interval. The design team planned to use an inherently reliable engine cycle and numerous state-of-the-art technologies derived from the SSME to fuse the knowledge and experience of the first-generation space shuttle program with second-generation research and technology development.

The COBRA engine was one of two hydrogen-fueled engine designs being evaluated for use as a first- or second-stage option for a next-generation reusable launch vehicle. Kerosene-fueled engines were also being considered for the first-stage booster. The engine’s key design features are shown in Table 4-5; a schematic of COBRA is shown in Figure 4-17.

TABLE 4-5 COBRA Engine Key Design Characteristics

Characteristics

Cobra

SSME (Block II)

Propellants

H2/O2

LOx/LH2

Thrust (lbf)

 

 

sea level

492,590

393,800 lbf at 104%

vacuum

600,000

448,800 lbf at 104%

Isp (sec)

 

 

sea level

373.3

363

vacuum

454.7

452

Chamber pressure (psia)

3,000

2,994

Mixture ratio

6.0

6.03:1

Engine T/W, at vacuum

75

73.12

Engine length (in.)

180

169

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-17 COBRA engine schematic. SOURCE: Pratt & Whitney.

XRS–2200 Engine, Boeing Rocketdyne

The aerospike engine (Figure 4-18) is a LOx/LH2 gas generator cycle engine. Each engine has a single oxidizer turbopump, a fuel turbopump, a gas generator, a combustion wave ignition system, 2 aerospike nozzle ramps, 10 thrust chambers (“thrusters”) per ramp, 2 redundant engine controller digital interface units, and associated plumbing, valves, and EMAs. Table 4-6 summarizes the XRS-2200 key engine characteristics.

TABLE 4-6 Key XRS-2200 Engine Characteristics

Characteristic

Value

Sea-level thrust (lbf)

206,800

Isp at 100% and MR 5.5 (sec)

332

Mixture ratio (MR)

4.5-6.0

Chamber pressure (psia)

830

Throttling (% of thrust)

57-100

Differential throttling (% of thrust)

±15

Dimensions (in.)

 

Forward end

134W x 90L

Aft end

42W x 90L

Forward to aft

90

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-18 XRS-2200, single-engine computer-aided design and manufacturing drawing. SOURCE: Boeing Rocketdyne.

The X-33 program, which began in July 1996, was a half-scale prototype of Lockheed Martin’s proposed single-stage-to-orbit (SSTO) concept, named the VentureStar. The program was set up as a unique cooperative agreement with Boeing Rocketdyne as the supplier of the XRS-2200 linear aerospike engine. Two of these engines were to be used to power the X-33 on suborbital flights to demonstrate the technology needed to proceed with the full-scale VentureStar.

A larger version of the XRL-2200 intended for the VentureStar vehicle was designated the RL-2200. It was designed for a sea-level thrust of 431,000 lb. Seven of these engines would have been used to lift the 2.2 million lb (GLOW) vehicle.

The philosophy of the X-33 program was to accept increased risk in order to achieve lower costs and quicker schedule. To do so, the XRS-2200 program relied heavily on the experience gained from Rocketdyne’s testing of a linear test bed engine from 1970 to 1972. Where possible, the X-33 program used existing hardware and/or designs. The turbopumps and the gas generator were based on J-2 and J-2S engines. Component testing was used for design development, proving margins, and qualifications. Software was tested with hardware-in-the-loop. Single engine testing on the Stennis Space Center’s A-1 test stand was used to verify the design. The two flight engines, in their dual-engine configuration, had had a short ignition test and were about to start acceptance testing when NASA decided not to renew its involvement in the cooperative agreement.

RD-0146 Engine, Pratt & Whitney

In addition to the RL-10 family of LOx/LH2 cryogenic upper-stage engines and the new RL-60, a higher thrust upper-stage engine currently under limited early development in-house, Pratt & Whitney

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

offers another cryogenic upper-stage engine, designed and developed by a highly experienced Russian maker of cryogenic engines. This engine, designated the RD-0146, is manufactured by Chemiautomatics Design Bureau (CADB) of Veronezh, Russia.

The engine produces 22,000 lb thrust at a minimum vacuum Isp of 451 sec. The RD-0146 is an expander cycle configuration with a wide operating range, including the capacity for many firings (restarts) in space. CADB also produces the RD-0120 cryogenic engine, which is the core booster engine for the Russian high-performance Energia launch vehicle, which was used to launch the Russian BURAN autonomous space shuttle. Pratt & Whitney is the distributor for this engine and claims that it stands behind the stated performance and operating characteristics as summarized in Table 4-7.

TABLE 4-7 RD-0146 Key Design Characteristics

Characteristics

Value

Thrust (lb)

22,000

Weight (lb)

534

Isp, vacuum (minimum) (sec)

451

Cycle

Full expander

Propellants

LH2/LOx

Mixture ratio

6.0

Restarts

Multiple

Furthermore, it claims that it will make the necessary modifications and certify the engine for any U.S. application and supply the production flight engines, as required.

With the recent about-face in NASA’s approach to its next-generation launch vehicle architecture, the four engines started under SLI remain essentially prototype concepts. It remains to be seen whether any of them or their derivatives will be attractive for a second-generation ORS. As stated in Recommendation 4-2, some of these large-engine concepts may be candidates for a long-term booster engine technology development program aimed at far-term replacement vehicles for Atlas and Delta.

TR-106 and TR-107 Engines, Northrop Grumman

Northrop Grumman has carried out the detailed design of a 1-million-lb-thrust booster rocket engine utilizing LOx/HC propellants as part of NGLT under NASA’s SLI. The authorization to proceed on this design was awarded in March 2003. The primary goal for the TR-107 engine program was to continue development of an engine that would increase the safety, reliability, and affordability of next-generation reusable space launch and transportation vehicles.

TR-106. Starting in the late 1990s, Northrop Grumman Space Technology (then TRW) undertook, using company funds, to design and build a large engine that could operate on either LOx/LH2 or LOx/RP-1. The engine was expected to replace solid propellant booster strap-ons with liquid propellant stages having on-command throttling shutdown and even restart. Liquid propellants were considered safer and more environmentally friendly. The engine also was envisioned for powering the first stage of expendable, or fly-back, boosters.

The engine designed and demonstrated in this effort was designated the TR-106. It had a planned sea-level thrust of 650,000 lb and was to be either pressure fed or operated with gas-generator-driven turbopumps in the propellant lines. The center pintle injector incorporated in the TR-106 engine can operate equally well using LOx with RP-1, ethanol, propane, methane, or LH2. This basic injector technology has a 40-year history of producing high-performance and totally stable combustion without baffles or quarter-wave acoustic chambers in engines with thrust ranging from 100 lb to 650,000 lb (Yang and Anderson, 1995; Yang et al., 2003).

The concept was originally developed at Space Technology Laboratories (STL) in 1960 as a 20:1 throttling injector for a 500-lb-thrust space-maneuvering thruster using dinitrogen tetroxide (N2O4) with

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

monopropellant hydrazine (N2H4). That throttling design was scaled up to 10,000 lb thrust with 10:1 throttling using N2O4 with N2H4 50-UDMH 50 (Aerozine-50) and was used in the lunar module descent engine (LEMDE) for the Apollo program. A fixed-thrust version of the engine was used on the second stage of Thor-Delta in the 1980s, where it accomplished more than 65 completely successful launches.

In the competition for the LEMDE between 1962 and 1964, STL explained to Northrop Grumman and NASA the reason for the engine’s bomb-demonstrated dynamic stability characteristics. The LEMDE combustor satisfied a fundamental design criterion: It did not provide significant energy sources in the antinode regions of a resonant chamber mode. Furthermore, it satisfied another criterion by putting almost all of the combustion energy source close to the nodal regions of any potentially destructive chamber mode. Satisfying this criterion assures the dynamic stabilization of initiating disruptions that always occur in a rocket combustion chamber.

For practical design reasons, and to throttle that engine with a single moving injector part while controlling the injector face heat transfer, the central injection criterion for stabilizing the first tangential mode (transverse or spinning) was best satisfied by a single coaxial pintle element. In fact for practical configuration and fluid dynamic reasons, such central element injection has also been demonstrated to dynamically stabilize the first radial mode.

The same basic pintle injector geometry has been tested at thrusts of 40,000 and 250,000 lb operating with N2O4/A-50; at 50,000 lb with LOx/RP-1; and at 40,000 and 650,000 lb with LOx/LH2. The basic central injector element scales essentially photographically—that is, for a given injector pressure drop, thrust is proportional to the square of pintle injector diameter, from 10,000 to 1 million lb thrust. The cutaway in Figure 4-19 is for a throttling high-thrust configuration and would look much the same at any thrust level.

FIGURE 4-19 Cutaway design of TR-106 engine. SOURCE: Northrop Grumman (TRW).

The 650,000 lb thrust TR-106 engine is shown in Figure 4-20. This injector technology has been demonstrated also with H2O2 and ethanol for storable propellant upper-stage and in-space applications. In 40 years of firings at every thrust level and with every propellant combination, there has never once been a case of combustion instability with this centrally located single-pintle injector.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-20 650,000-lb thrust TR-106 engine. SOURCE: Northrop Grumman.

TR-107. As stated earlier, a primary goal for NASA’s SLI contract for the TR-107 program was to continue development of an engine that could increase the safety, reliability, and affordability of next-generation reusable space launch and transportation vehicles. The contract specified that a high-pressure oxygen-rich staged-combustion (ORSC) cycle was to be used with the propellants LOx/RP-1.

In its earliest concept phase, the TR-107 had a central pintle injector for both the main combustion chamber and the LOx-rich preburner. However, performance and risk analyses soon indicated that the main injector should switch to a distributed coaxial multielement design, given that a single ORSC preburner would be used to drive both the fuel and the oxidizer turbopumps. The oxygen-rich preburner (basically a gas generator) retained the pintle injector because its size was within the pintle injector LOx/RP-1 test database and because Northrop Grumman wanted to retain the inherent stability of the pintle injector. It also wanted to give the preburner throttling capability for mission flexibility and to allow future growth.

The TR-107 engine was one of several SLI candidate engines (some of which were described above) that could be used to provide primary propulsion for the Earth-to-orbit (ETO) stage of future reusable launch vehicles. Several primary technology objectives of the TR-107 program were accomplished before the SLI efforts were terminated:

  • Successful demonstration of a duct-cooled chamber, which eliminates the need for conventional cooling channels;

  • Successful demonstration of the preburner pintle injector and propellant mixing, which stabilizes throttling and enables slow, controlled-start-up transient performance;

  • Specification of properties for materials compatible with oxygen-rich environments, which eliminates the need for additional coatings and liners;

  • Incorporation of mature combustion devices that minimize parts count for greater reliability and operability; and

  • Systems-engineered design optimization to minimize cycle pressures, which provides margin to increase engine life. Additional design details of the engine are shown in Figure 4-21.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-21 Details of the TR-107 engine. SOURCE: NASA, Marshall Space Flight Center.

The central pintle injector technology is also used in the engines for the FALCON program by SpaceX (with LOx/RP-1) and by AirLaunch (with LOx/propane). It has been used in every Northrop Grumman (TRW) in-space bipropellant maneuvering engine and in the alternate current thrusters on dozens of major satellites. Because it can operate with nearly any propellant combination, engines incorporating it are excellent candidates for either the reusable booster or second stage of ARES or future ORS using simple gas-generator, driven propellant-line pumps.

Fastrac and MC-1 Engine, NASA

In the late 1990s, an in house rocket engine design and development project was initiated at NASA MSFC to give the younger propulsion engineers some real hands-on hardware experience. The objectives of this project were ultimately to design, build, test, and evaluate a 60,000-lbf LOx/RP-1 low-cost, full-up rocket engine (prototype-level) to be demonstrated in a test bed. Senior MSFC management believed that a number of the capable young rocket propulsion engineers who had joined the MSFC team had never, in a decade, participated in the development of a rocket engine or even seen one. It was felt that a test bed/prototype engine design and hardware project, conducted entirely inhouse, would be very effective in training and preparing these young engineers, under the guidance of a few experienced rocket development engineers at MSFC, to participate in future NASA rocket development and flight hardware procurement.

The project was first called the Fastrac engine and conducted according to all the NASA design and development ground rules and project criteria. However, just before the formal Fastrac preliminary design review, the engine requirements were changed in a major way so that it could be applied to the reusable X-34 of the Pathfinder/X-vehicle demonstrator program, and the engine was renamed the MC-1. Eventually, the X-34 flight demonstrator program was cancelled by NASA, and all further work on the Fastrac/MC-1 engine was terminated. The X-34 was cancelled owing to NASA’s decision to terminate all activities associated with SSTO next-generation launch vehicles. The Air Force subsequently declined to pick up or fund any follow-on work.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

Some good technology was developed on this Fastrac engine project, the most notable of which was the low-cost Barber Nichols fuel and oxidizer pump assembly. These pumps were well along in their development and readiness for incorporation in the flight MC-1 engine. When the in-house MSFC project was terminated, the turbopump assembly elements were eventually shifted for use on two current new, small, low-cost, launch vehicle development projects jointly sponsored by DARPA, the Air Force Space and Missile Systems Center, and NASA. Those two projects are part of the FALCON small launch vehicle program.

Barber Nichols TPA elements are currently being used by the SpaceX version of FALCON to pump propellants to its 75,000 lbf LOx/RP-1 Merlin first-stage liquid booster engine (operating at 850 psia) and by Lockheed Martin Michoud to pump LOx to the first and second stages of the hybrid rocket motor for its version of the FALCON launch vehicle. Some of the background and results of the original NASA MSFC Fastrac turbomachinery development program are summarized next.

The original goal of the Fastrac turbomachinery program was to reduce the unit cost and the lead time required to produce turbomachinery for rocket engines. The major vendors involved with the fabrication of hardware for the turbomachinery are summarized in Table 4-8.

TABLE 4-8 TPA Major Vendors

Vendor

Responsibility

Summa

Engine prime contractor

Barber Nichols

TPA pump subcontractor

Howmet Inc.

Investment castings

EG&G

Bellows seals

Wollaston Alloys, Inc.

Sand castings

Walcolmanoy

Brazing

Barden

Bearings

As program requirements evolved for the Fastrac engine, thrust was increased to 24,000 lb and, finally, to 60,000 lb. At the 60,000-lb thrust level, the target first unit cost for the required turbomachinery was $300,000, with a target production unit cost of $150,000. To achieve lower cost and faster acquisition time, the critical design team pursued a reduced-part-count design. As part of this approach it decided that the oxidizer and the fuel pumps would be placed on a common shaft and driven by a single turbine. The basic performance requirements for the TPA are listed in Table 4-9.

TABLE 4-9 Basic Performance Requirements for the Fastrac/MC-1 Barber Nichols TPA

Part

Specification

Oxidizer pump

 

Fluid

LOx

Inlet pressure (psia)

46.0

Discharge pressure (psia)

919.0

Mass flow rate (lbm/sec)

138.61

Fuel pump

 

Fluid

RP-1

Inlet pressure (psia)

28.0

Discharge pressure (psia)

959

Mass flow rate (lbm/sec)

63.96

Turbine

 

Fluid

Combustion products

Speed (rpm)

20,014

Inlet pressure (psia)

550

Discharge pressure (psia)

65

Inlet temperature (°R)

1600

Mass flow rate (lbm/sec)

7.1

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

The turbopump for the MC-1 engine is a common shaft design with the oxidizer pump located forward, the fuel pump amid shaft, and the turbine in the aft. The oxidizer and the fuel pumps comprise an axial flow inducer followed by a radial flow impeller. The oxidizer pump has an axial inlet, while the fuel pump has a plenum inlet fed by two inlets 180 apart. The turbine section is composed of a single-stage transonic impulse stage with a bladed disk turbine wheel followed by a set of exit guide vanes. The turbine has a plenum inlet and an axial discharge. Rolling element bearings support the pump rotor.

This arrangement allowed the elimination of a turbine wheel, a turbine housing, and hot gas ducting between turbines. Engine system benefits of this design include elimination of support brackets for an additional turbopump assembly, a requirement for only one turbine discharge duct, and reduced potential for operational runaway of either pump. Along with the benefits came several issues that had to be addressed. The most notable of these were the compromise in shaft speed needed to place both pumps on the same shaft driven by a single turbine, the design of the inlet for the fuel pump, and thermal conditioning of the TPA prior to engine start.

After the decision was made to have a common shaft, the arrangement of the elements on the shaft had to be set. The LOx pump was placed on one end of the shaft and given an axial inlet. This was done because the LOx pump had more stringent suction performance requirements, and this arrangement optimized the ability to meet those requirements. If the fuel pump had been placed on the opposite end of the shaft, it also could have had an axial inlet; however, maintaining the concentricity of the housings with the turbine in the middle was a major concern. The pumping elements were placed back to back to assist in managing the axial thrust in the TPA. The plenum inlet and axial discharge to the turbine were chosen to assist with engine packaging. The gas generator is connected directly to the inlet of the turbine inlet manifold and the turbine discharge is routed down the side of the engine’s nozzle.

Most of the TPA components are fabricated from Inconel 718. Housings are conventional vacuum investment castings. A cross section of the Barber Nichols TPA is shown in Figure 4-22. Modifications of this basic turbopump design are incorporated as line pumps on one of the FALCON small vehicles.

FIGURE 4-22 MC-1 Barber Nichols TPA. SOURCE: NASA, MSFC.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×
Materials and Chamber Cooling, Pratt & Whitney

Pratt & Whitney is working on advanced high-temperature aluminum alloys. Initial characterization indicates their specific strength is 2.5 times as great as that of current steel jackets at elevated temperatures. The company is investigating structurally compliant chamber walls to accommodate increasing pressures and thermal stress and to supply the required cooling characteristics.20

Sequential Feed System, University of Alabama in Huntsville

Current launch vehicles, missiles, and high-performance upper stages using liquid propellants rely on either pressure-fed or turbopump systems. As shown in Figure 4-23, in-space liquid rocket engines with relatively low total impulse, low pressure, and low thrust typically use conventional pressure-fed systems.

FIGURE 4-23 Representative regions of application for pump-fed, sequential feed, and turbopump systems. SOURCE: Department of Mechanical and Aerospace Engineering, University of Alabama in Huntsville.

Conventional pressure-fed systems are low cost, relatively heavy, and very reliable. They typically deliver relatively low payload mass. Systems with higher total impulse currently use turbopumps, which are relatively light but expensive and which to date are not available for low-thrust (1,000-5,000 lb), in-space systems or for moderately sized missiles. Turbopumps typically provide a relatively high delivered payload. An alternative approach, the sequential feed system (SFS),21 offers a means of reducing the cost and weight and improving the reliability and performance of launch vehicles, missiles, and upper-stage and in-space propulsion systems. McDonnell Douglas, which originally conceived and patented the SFS, donated its rights to the SFS concept to the University of Alabama in Huntsville (UAH), which has developed a system sizing code, a full-scale SFS test bed, and the design for an advanced version of the

20

Notes taken from site visit to Pratt & Whitney Space Propulsion, West Palm Beach, Florida, on May 18, 2005, by committee members Yvonne Brill and D. Brian Landrum.

21

Also referred to as reciprocating feed systems (RFS).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

SFS, including integrated valves. Results of these studies show that the weight of the SFS-delivered payload rivals, and can exceed, that of turbopump systems. With its failsafe capability, the SFS offers an approach deserving of further consideration.

The basic concept relies on relatively small conventional propellant tanks and valves that are sequentially filled and pressured to a relatively high pressure. The propellant is then expelled to the engine and the small tanks are vented and refilled from large main tanks, which are operated at low pressure. Typically, the system uses three tanks but has an intrinsic failsafe capability in that it can operate with just two tanks. Steady, controlled flow and rapid, deep throttling have been demonstrated with flow rates commensurate with 20,000-lb-thrust engines in a test bed built as a collaborative effort with NASA’s MSFC. Further studies are continuing. The SFS uses relatively low-mass, low-pressure main tanks for the fuel and oxidizer, connected with two or three small, high-pressure fuel and oxidizer tanks. The small tanks alternately expel propellant, vent, and are then refilled and pressurized in sequence, to maintain steady and/or modulated high-pressure flow to the engine. The SFS relies on this sequential process to provide liquids at high pressure; there is no mechanical pump hardware involved, since the SFS comprises valves and small tanks. Analysis to date shows that for low pressures and low total impulse, the conventional pressure-fed systems would be used. For moderate to high engine pressures, and for high total impulse (with attendant large propellant masses and large tanks), then both SFS and turbopump systems would be preferred. Only with very high chamber pressures would the turbopump system offer a weight savings relative to the SFS. Other benefits include built-in redundancy and fail-operational modes, which offer improved system reliability and improved propellant management in a microgravity environment.

NASA and several companies have approached UAH about using the SFS on launch vehicle concepts to address affordable responsive spacelift (ARES) requirements (Blackmon and Eddleman, 2005).

Apparently Superior Foreign Technologies

The committee looked at foreign (Russian, Chinese, and Indian) rocket and launch vehicle technologies to determine whether some of them were significantly superior to available U.S. technologies. The apparent growth in Chinese capabilities appears to be tied partially to the transfer of technology from Russia. In addition the Russians have sold an LOx/LH2 upper stage to the Indian program, which could significantly increase its capability. Reverse technology transfers, however, have significant limitations.

Foreign booster engines include the NK-33 and the RD-180, provided through Pratt & Whitney. The NK-33 inventory has been completely acquired by Aerojet. Pratt & Whitney is in the process of completing technology transfer from NPO Energomash with a goal to co-produce the engine by 2010-2012. Although both parties have signed the transfer contracts, there have been cases where these contracts have been terminated prior to total transfer.

The only foreign upper-stage engine identified is the Russian RD-0146, made by CADB. Pratt & Whitney has agreements for procurement and possible production of this engine in the United States. The performance of this LOx/LH2 expander cycle engine is similar to that of the RL-10.22

From these sources it does not appear that foreign launch vehicles are using propulsion technologies that represent state-of-the-art advances over technologies used by the United States. Both the Chinese and Indian contenders have bustling launch vehicle programs, and the Russian engineering influence is clearly seen in the product. The United States is already aware of the performance advantages offered by some rockets from the former Soviet Union, and U.S. companies like Boeing (whose Sea Launch, for example,

22

Aviation Week’s annual aerospace source book gives a comprehensive listing of current U.S. and foreign launch vehicles. More in-depth data on these launchers is also available in International Reference Guide to Space Launch Systems, published by the American Institute of Aeronautics and Astronautics.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

used the Zenit), International Launch Services (Proton), and Lockheed Martin (its Atlas V uses RD-180) have incorporated these engines.

DEFINING DOD AND AIR FORCE NEEDS FOR PROPULSION TECHNOLOGIES AND TOOLS

Systems Engineering

Meeting the objectives of ARES and ORS vehicles for first, second, and—conceivably—third stages may necessitate a number of new rocket propulsion subsystem technologies in addition to those that exist in applicable qualified subsystems or that are embodied in the designs of the new conceptual engines. A rigorously disciplined, integrated total systems engineering process is required from the very start in order to outline the total mission trade space and to select propulsion system requirements for each stage of each total vehicle system concept. This engineering process dictates that the primary criterion for defining these requirements be “mission success.” Because overall mission success is critically dependent on all of the engineering design considerations discussed below, the process must have continuous feedback from each of these design activities as they are evolved.

First is the definition of rocket propulsion concepts and technologies capable of meeting those tiered-down requirements. Crucial to an integrated total systems engineering process for assuring success is forcing the explicit identification of the design criteria validation status of all proposed critical technologies. This identification is the dominant factor in making objective evaluations across a broad propulsion systems trade space of development engineering schedule and cost risk and of subsequent propulsion systems’ operational and life-cycle cost risks.

Identifying the unvalidated design criteria associated with all propulsion systems concepts proposed to meet Air Force new ARES and ORS vehicle needs would define high-priority propulsion technologies programs requiring immediate DoD/Air Force investments. Such technologies are not likely to be transformational, given the existing ORS roadmap, but they are absolutely crucial to the success of ARES/ORS.

The committee’s rocket panel has identified two technology areas as important tools and elements of the design criteria database.

Integrated Totals Systems Engineering Process

Because it can determine how the DoD/Air Force technology mix should be restructured to effectively support ORS, one of the highly-leveraged critical technologies requiring immediate effort is the further evolution of an integrated total systems engineering process, with mission success as the primary selection criterion (See Recommendation 4-1).

Virtual Rocket Engine Design and Testing

Funding the continued evolution of computerized high-fidelity virtual designs for engine and propulsion systems that can be tested on a virtual test stand and ultimately in virtual flight could be one of the most cost-effective highly leveraged investments that DoD and the Air Force ever make. The potential for reducing the huge costs associated with the usual cut-and-dried development of rocket engines and their propellant systems is great when the time line is extended into the indefinite future.

According to a paper to be published in the fall of 2006, advanced software simulation tools have helped to reduce cost and development time for new propulsion systems by allowing the design team to perform dynamic system analysis before the hardware is fabricated (Sackheim, 2006). These tools provide valuable design insight and can speed system optimization, moving it forward to the early phases of the development program. An optimized system normally results in a simpler, more reliable design

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

with significantly lower development costs. The ability to adequately model hardware and systems also speeds the development process and reduces the number of test-fail-fix-test cycles (see Box 4-1).

Box 4-1

Virtual Testing as an Enabler for the RS-68 Rocket Engine

The J-2 engine, developed in the mid-1960s, was used on the second and third stages of the Saturn V rocket. During development of the RS-68 engine, currently used in the Delta IV common booster core, scaling data from the base J-2 engine in combination with modeling and simulation tools were used to reduce duration of the test-fail-fix-test cycles by approximately one-half (Wood, 2002). This approach was also used to reduce development costs for the SSME Block-II engine and is being used in the development of the J-2X engine planned for use on the NASA crew launch vehicle upper stage.

This observation on the cost-driving characteristics of the hardware applies to recurring operational costs for expendable launch vehicles as well. The average recurring cost breakdown for the Atlas, Centaur, Delta II, Titan II, and Titan IV launch vehicles is approximately 71 percent for the vehicle hardware. Of that hardware costs approximately 54 percent is driven by propulsion elements (engines, strap-on boosters, etc.), not including the tanks, which are classified as part of the vehicle structure for bookkeeping purposes. So here again, a concerted effort to reduce engine costs will also greatly reduce the recurring and developmental costs of the expendable launch vehicle. Furthermore, it can be shown that engine costs are mostly a function of operating chamber pressure and parts count, both of which basically are a strong function of the type of power cycle (i.e., pressure-fed, gas generator, expander, or staged combustion).

From the above key points, it should be readily apparent that simplified engine designs and existing databases, in combination with the extremely advanced analysis and modeling tools based on the advanced CFD codes now available, should go a long way toward reducing the developmental and recurring costs of booster and upper-stage engines for future expendable heavy-lift launchers and in-space and descent/ascent engines and systems for human spacecraft (Sackheim, 2006).

Modeling and Simulation

There are three levels of modeling and simulation (M&S) tools: system, engineering, and research.

  • System level. Although system-level tools model a broad range of primary and subcomponent systems and are relatively fast, they tend to have low order accuracy and use empirically based global design relations (e.g., total engine mass based on chamber pressure). Engine design inputs are typically parameters such as specific impulse and T/W. Many of the existing codes incorporate models for cost and reliability. However, because these models have generally been validated for engines and launch vehicles that are similar to existing configurations, they may produce inaccurate results for new systems under consideration. Performance, cost, and reliability data are especially lacking for reusable rocket engines and launch vehicles of interest to the Air Force.

  • Engineering level. Engineering-level tools incorporate a mix of analytical techniques with empirical corrections (fudge factors). These models are typically one-dimensional but relatively fast. They often rely on databases, especially for performance, cost, and reliability, and many of these models are proprietary, with databases specific to a particular company’s engines and launch vehicles.

  • Research level. Research codes are primarily used in academic and government research labs. These codes include fundamental physical models that provide more accurate results than empirical-based engineering codes. However, research codes are often focused on isolated, or

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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uncoupled, phenomena (combustion, fluid dynamics, finite-element models, etc.). They typically require a significant level of user experience in grid generation, computational convergence, and the like. They may require significant computer resources such as supercomputers or large, distributed computer architectures (clusters).

Current Status of Air Force M&S Tools

Current Air Force M&S codes and environments are typically 30-40 years old, empirically based, and often use one-dimensional analyses. Many of them require extensive experience (knowledge of nuances and rules of thumb) to effectively run. The various discipline codes (fluids, thermal, finite element models, etc.) are generally stand-alone, with no integration or interfaces. The bottom line is that Air Force M&S tools are aging and are limited in their ability to support an integrated design process. M&S tools for air-breathing vehicles appear to be well ahead of those for rockets and launch vehicles. A significant upgrade effort in the rocket area is needed. This is becoming even more critical as the aerospace workforce ages. New M&S tools must capture the technical expertise of rocket design for future novice engineers (Huggins, 2005).

Air Force-Funded M&S Initiatives

Under IHPRPT, AFRL is leading a baseline study using M&S tools to evolve a current three-stage solid rocket launch vehicle into an advanced two-stage vehicle. Optimizing of the trajectory (range) drives the design. A solid rocket motor spreadsheet was used to investigate the effects of propellant density on Isp. The study also used the Integrated Solid Modeling Analysis Tool and the Integrated Propulsion Analysis Tool for liquid engine performance estimates. The effort included enhancing the capabilities of the POST2 trajectory code and allowing it to run on a personal computer (Hilton, 2005).

As discussed earlier, AFRL is managing an IHPRPT/USET program led by Aerojet and Northrop Grumman to improve the computational design and analysis tools used for advanced rocket engine development. This includes improving interconnectivity and efficiency while providing optimization capabilities. The effort is using commercial/university software programs and is upgrading existing codes in selected areas. One of these areas is turbopump analysis and design. A small business, C&R Technologies, is upgrading software to analyze secondary flows from a thermal/fluid perspective. This includes labyrinth seals, hydrostatic and other bearings, and system-level performance aspects such as torque, pressure drop, choking, dissipation, axial thrust, and radial torque for start-up events (Hilton, 2005).

NASA P-STAR Code

The bulk of NASA systems analysis work until now has been focused at the vehicle level. Only top-level propulsion requirements have been addressed through parameters such as T/W and Isp. NASA has developed several comprehensive tools (Generic, ROCETS, etc.), but they are not flexible enough to allow rapid engine trade studies. The Propulsion Sizing, Thermal, Accountability, and Weight Relationship Model (P-STAR) is being developed and used by NASA’s Space Transportation Directorate at MSFC. P-STAR is a flexible and scalable propulsion system design model that provides first-order engine balance, thermal balance, reliability/safety assessment, credible weight estimate, and cost prediction. The P-STAR physics-based environment includes approximately 70 Excel worksheets with add-ins, virtual basic code, and data tables that are subject to International Traffic in Arms Regulations and/or contractor proprietary rules (Leahy, 2005).

For rocket engine analysis and design, the code uses a bottom-up approach. The user specifies the thrust, chamber pressure, nozzle area ratio, and mixture ratio. Models such as the Chemical Equilibrium Analysis code are then used to compute engine performance parameters such as Isp. P-STAR includes a library of engine cycles for boosters, upper stages, and orbital maneuvering systems (Leahy, 2005).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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NASA/University of Alabama in Huntsville

Overall launch systems development cost can be substantially minimized by optimizing propulsion system components using historical engine data, a propulsion thermochemical code, and optimization tools and techniques. A joint effort between the University of Alabama in Huntsville (UAH) and NASA MSFC is attempting to fully integrate key propulsion variables such as Isp, engine mass, gross liftoff weight (GLOW), nozzle area ratio (AR), chamber pressure (PC), initial T/W (Twi), and oxidizer/fuel (OF) ratio into a model for closed-loop analysis of launch vehicle concepts. The model process flow has been demonstrated for a SSTO system using liquid oxygen (LOx) and LH2. Generic algorithm solvers are used for concept optimization (Shelton et al., 2005).

Figure 4-24 illustrates the calculation of higher level engine performance parameters by using PC, AR, OF ratio, and the Cequel thermochemical code. This methodology has not been used previously to design and optimize a closed-loop system.

FIGURE 4-24 Engine design using fundamental parameters and thermochemical codes. SOURCE: Shelton et al. (2005).

The Cequel output and input structure designed in the linkages proved that a true closed-loop system produces the results in a much more efficient way. The alternative to integrating thermochemical results into the model is to use databases and extrapolate the results for a given input condition. A key aspect of the model was which of the two methods was used—parameters of the system or propulsion variables produced in the design process—to determine engine mass (Shelton et al., 2005).

NASA Constellation University Institutes Project

The main goal of the Constellation University Institutes Project (CUIP) is to provide the NASA Constellation Systems within the Exploration Systems Mission Directorate with the products of long-term

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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research and development.23 The original three CUIP groups were the Institute for Future Space Transport, led by the University of Florida; the Space Vehicle Technology Institute (SVTI), led by the University of Maryland; and the Rocket Engine Advancement Program (REAP2), led by UAH. As of August 2006, the CUIP consortium consists of 18 universities, led by the Johns Hopkins University Applied Physics Lab, managed by NASA, and advised by a board that includes the Air Force Office of Scientific Research (AFOSR), AFRL, and the major commercial engine and launcher manufacturers. The member universities are divided into five virtual institutes based on the following technical areas: thrust chamber assembly, propellant storage and delivery, structures and materials for extreme environments, reentry aerothermodynamics, and systems analysis. The member universities currently receive approximately $8 million dollars of NASA funds.

One of the CUIP goals is to enable the use of CFD and computational structural design tools for multidisciplinary simulation and design of rocket engines, including the performance, life, and stability of the thrust chamber assembly and the supporting infrastructure. This includes increasing the fidelity, robustness, and accuracy of these tools and implementing a rigorous verification and validation process (AIAA, 1998). Different levels of simulation are addressed starting with the complete engine system and then decomposing it into subsystems (e.g., injector faceplate), model problems (e.g., a single injector element), and unit problems (e.g., mixing). Increasingly complex benchmarking experiments are performed to provide experimental data for the simulation readiness level of the computational codes and physical submodels.

FIGURE 4-25 Calculation of conjugate gradient heat transfer for supercritical hydrogen in a high-aspect-ratio copper cooling tube. SOURCE: University of Alabama in Huntsville.

An example of a CUIP problem with a model, shown in Figure 4-25, involves calculating conjugate gradient heat transfer for supercritical hydrogen in a high-aspect-ratio copper cooling tube. A companion experiment has been developed to provide benchmark data.

23

For more information on CUIP, see its official Web site at http://microgravity.grc.nasa.gov/Exploration/external/cuip_about.htm. Last accessed on August 8, 2006.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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University of Illinois at Urbana-Champaign Center for Simulation of Advanced Rockets

The U.S. Department of Energy (DOE) Accelerated Strategic Computing Initiative/Academic Strategic Alliances Program encouraged the University of Illinois at Urbana-Champaign to establish the Center for Simulation of Advanced Rockets (CSAR) in 1997.24 The goal of CSAR is the detailed, whole-system simulation of solid propellant rockets from first principles under both normal and abnormal operating conditions. The design of solid propellant rockets is a sophisticated technological problem requiring expertise in diverse subdisciplines, including the ignition and combustion of composite energetic materials; the solid mechanics of the propellant, case, insulation, and nozzle; the fluid dynamics of the interior flow and exhaust plume; the aging and damage of components; and the analysis of various potential failure modes. These problems are characterized by very high energy densities, extremely diverse length and time scales, complex interfaces, and reactive, turbulent, and multiphase flows.

CSAR is composed of nine research teams to address the specific needs of each aspect of the simulation. The team responsibilities are (1) combustion modeling and corresponding codes for simulating the burning of composite propellant and the thermo-mechanical behavior of energetic materials; (2) hydrodynamical modeling and corresponding codes for simulating the interior cavity flow and exhaust plume; (3) solid-mechanical and thermal modeling and corresponding codes for simulating the case, nozzle, insulation, and propellant; (4) performance evaluation and tuning of individual component codes as well as the integrated system code; (5) parallel numerical algorithms and algorithms for mesh generation and adaptive refinement; (6) identifying test problems for system and component code verification and validation; (7) physical coupling and time stepping; (8) software integration to define software and data interfaces for coupling component codes; and (9) modeling and corresponding codes for assessing various failure modes and the effects of aging and damage on constituent materials.

Pratt & Whitney

Pratt & Whitney Space Propulsion has been updating its collaborative analysis and design processes for complex systems, with a focus on propulsion and power for use in an integrated concurrent engineering analysis and design approach. This integrated total aerospace power system (ITAPS) environment focuses on the formulation and validation of methodologies and processes that create a responsive and rapid advanced design and analysis environment for evaluating aerospace systems. Based on experience in the synthesis of all subsystems on military and commercial aircraft, ITAPS uses an integrated approach to examining system architectures, mission requirements, and the interaction of all the elements of an aerospace system (i.e., propulsion, power, electrical systems, vehicle sizing, and vehicle flight performance) (Joyner and McGinnis, 2004). The ITAPS approach is based on the integration of multidisciplinary design analysis tools to evaluate all manner of aerospace vehicle systems. These systems include lower-level coupled subsystems (e.g., the engine health management system, power and power management systems, and the propulsion system) and higher-level total systems like expendable and reusable launch systems with all their subsystem elements. The design integration of these “systems of systems” is examined at a higher system level using Phoenix Integration’s ModelCenter software.

24

Detailed information on this center and its simulation efforts is available at http://www.csar.uiuc.edu/F_info/AboutCSAR.htm#Intro. Last accessed on August 28, 2006.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

FIGURE 4-26 ITAPS functional analysis program. SOURCE: Joyner and McGinnis (2004).

As shown in Figure 4-26, subsystem functional elements are determined and then the functional interdependencies are identified. The functional description at the system level is used to define the functional modeling requirements that make up an ITAPS integrated model. Detailed models for subsystems can be easily integrated into the simulation architecture. These models can be incorporated as spreadsheets, empirical engineering codes, or structured CFD codes. This allows rapid synthesis and evaluation of the functional attributes of an aerospace system (Joyner and McGinnis, 2004).

Lockheed Martin Space Systems

With independent R&D funds and through their participation in the DARPA FALCON program, Lockheed Martin Space Systems Company has developed a hybrid rocket design tool. The tool has been validated through the testing of various scale boosters and calibrated for multiple port configurations.

Finding 4-10. The Air Force and NASA efforts to go from empirical-based codes to physics-based codes are very important. Without disputing the niche applications for empirical-based codes, physics-based codes have broader ranges of applicability and generally give more accurate results. Using commercially available and proven codes makes the results gotten by any entity more easily understood and comparable with other results, since more people are familiar with the capabilities and limitations of commercial codes. Also, the commercial suppliers bear most of the cost of keeping their codes up to date and compatible with the latest computational capabilities. This is a very important enabling technology.

The Air Force needs conceptual vehicle design tools to conduct honest broker assessments of the system benefits of propulsion technology in a timely manner and to evaluate concepts proposed by industry. This includes M&S tools that facilitate comparative trade-offs of potential propulsion concepts (e.g., cycles and fuels) and component technologies (e.g., turbopumps vs. pressure-fed). Ability to design and analyze new virtual engine concepts is also needed. The tools would benefit from being multidisciplinary (capable of being used for coupling fluids, structures, cost, and reliability, for instance).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

The growing capabilities and declining costs of computer resources mean it is now possible to incorporate higher fidelity, physics-based analysis codes into a total architecture systems engineering approach. However, the models must be validated, necessitating investment in tests of subcomponents and rocket engines in a range of physically correct operating environments. As the workforce ages, it is necessary to capture their technical expertise for future generations of engineers.


Recommendation 4-10. Because the development of the required quantitative systems engineering and design tools requires a national team approach, the Air Force and DoD should provide the leadership to assure that this critical capability becomes the very best in the world. DoD should provide a significant fraction of the resources necessary to accomplish this goal. As the primary members of the national team, the Air Force and DoD should provide mission definitions and system requirements to interactively identify and prioritize tool capability requirements. The Air Force should establish a process for maintaining and upgrading modeling and simulation tools. Commercial code developers (e.g., FLUENT, ANSYS) can provide interfaces and proven algorithms. Universities and government research laboratories can provide new models incorporating fundamental physics (e.g., large eddy simulation). Finally, industry can provide cost, operations, and practical subsystems models. One possible approach is to establish a consortium consisting of universities, companies, and academic institutions under Air Force leadership with an approach similar to the National Project for Applications-Oriented Research in CFD Alliance.25 The resultant world-class enhanced modeling and simulation capabilities would provide the Air Force and DoD with a transformational process for objectively identifying and prioritizing their research and technology investments.

Rocket Engine and Motor Test Beds

Flexible, responsive test beds for large engines and solid-propellant motors are needed to enable ORS success. In today’s market environment, rocket propulsion contractors can no longer support large engine testing facilities and test beds. Limited funding is one of the main reasons that this is not being done at the present time.

Important Technologies for Propulsion Systems

System Health Monitoring

If the Air Force intends to develop reusable rocket engines for ORS, then it is imperative that vehicle health management technologies be developed to monitor propulsion systems in flight, determine if problems are in the making, and, hopefully, have on hand options to resolve them before they become catastrophic events. All health status data must be provided to the turnaround launch site if fast rework and turnaround times are to be achieved.

Liquid Propellant Rocket Engines

DoD and the Air Force need to focus on technologies for propulsion systems that use liquid propellant. Other than the IPD currently being tested under NASA limited funding at Stennis Space Center, DoD is not funding any technology development that looks at, for example, new pumps, turbines, advanced power cycles such as oxidizer-rich staged combustion, single preburners, or expander cycles operating at higher pressure and with better performance than the RL-10. Although a number of detailed designs for these elements have been incorporated into conceptual engines, the designs have not been validated and they must therefore be considered as too risky to be committed to ARES. Where to focus DoD funding should be determined by the results of ARES systems engineering. The propellant needs to

25

For additional information on NPARC, please see http://www.arnold.af.mil/nparc/. Last accessed on August 28, 2006.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×
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

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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).

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
×

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.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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OSC. 2004. Minotaur Users Guide. October. Available online at http://www.orbital.com/NewsInfo/Publications/Minotaur_Guide.pdf. Last accessed on March 30, 2006.

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Weeks, David, Steven H. Walker, and Robert L. Sackheim. 2005. Small satellites and the DARPA/Air Force FALCON program. Acta Astronautica: AA2399.

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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.

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Scott Smith. “Air-based vertical launch,” Presentation to committee member Gerard Elverum on December 19, 2005.

Suggested Citation:"4 Rocket Propulsion Systems for Access to Space." National Research Council. 2006. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs. Washington, DC: The National Academies Press. doi: 10.17226/11780.
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Rocket and air-breathing propulsion systems are the foundation on which planning for future aerospace systems rests. A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs assesses the existing technical base in these areas and examines the future Air Force capabilities the base will be expected to support. This report also defines gaps and recommends where future warfighter capabilities not yet fully defined could be met by current science and technology development plans.

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