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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs 5 Rocket Propulsion Systems for In-Space Operations and Missiles INTRODUCTION Current and future strategic and warfighter needs for satellites and in-space vehicles exist in the following areas: Strategic assets for communication, early warning, Earth observation, navigation, reconnaissance, surveillance, and weather; Technology development work in space; and Responsive space operations. All DoD and other U.S. strategic satellites and technology platforms require propulsion subsystems operating in space to provide the impulse necessary to adjust velocity, change orbit altitude, and provide attitude control, station keeping, and end-of-life deorbit. These propulsion needs are being satisfied currently by state-of-the-art chemical propulsion, and, increasingly, by electric propulsion subsystems. It should be noted here that the state of the art has been undergoing major changes over the past 15 years and therefore represents a very advanced capability in many areas. For the kilogram-weight-class satellites with unique mission capabilities that are being developed, micropropulsion systems may be all that is required for maneuvers other than rapid inclination change. However, in contrast to rocket propulsion for access to space and near space, the range of potential improvements for in-space propulsion thruster performance and for electric power generation and energy storage is still very large. Some of these technologies, such as various types of electrically powered thrusters or high-energy monopropellants, have the potential to be transformational for in-space military systems capabilities. Air Force and DoD long-range plans have identified some needs and are still working out other needs for many types of operational maneuvers in space and near space. The systems architecture for seamless air-space operations has been termed operationally responsive spacelift (ORS). The elements of the ORS architecture are depicted in a very general form in Figure 5-1. As indicated in Figure 5-1, the Air Force will continue to identify needs for responsive and rapid introduction or repositioning of military satellites or space vehicles of various types for surveillance, defensive or offensive deployment, access to local theater military operations, and, secondarily, for situational awareness. There are a number of approaches to meeting these various military needs. For large strategic and capital assets, one could utilize an onboard, low-thrust, highly fuel-efficient system such as a Hall-effect thruster that would fire continuously to complete a large station change or a repositioning maneuver at high specific impulse (Isp) (2,000 to 3,000 sec). However, large assets accomplish such maneuvers with velocity changes measured in feet per second per hour and would take weeks to travel, say, 10,000 mi.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-1 Responsive space utility. SOURCE: James (2005). A second approach would be to use a moderate-thrust (100-200 lb), modest-performance chemical propulsion thruster at 350-360 sec Isp, such as liquid oxygen (LOx)/monopropellant hydrazine using a cryocooler to keep the LOx from boiling away. Velocity changes of hundreds of feet per second can be achieved in minutes to hours, depending on the propellant mass available, for platforms of thousands of pounds, permitting position changes of thousands of miles per day. With higher thrust levels, maneuvering times could be quite rapid. One important technology that would permit multiple maneuvers of critical assets would be an on-orbit refueling system to resupply the propellants while changing stations. The on-orbit refueling capability would enable the space asset to stay alive as long as everything kept working and to make as many rapid station changes as required. The capability for on-orbit docking and refueling will be demonstrated with hydrazine and high-pressure helium in space by the end of 2006. A third approach to implementing rapid station changes would be to have a large (perhaps refuelable on orbit) space tug with high-performance electric propulsion for slow strategic moves or a high-thrust, modest-performance chemical system for responsive maneuvers. Or, a tug could have some combination of propulsion systems on board that would allow it to fly up, dock with a key space asset, and move it to the desired new operational location. The tug could then de-mate from the spacecraft and fly on to reposition other assets as needed. This approach would entail a small fleet of permanently based space maneuvering tugs that would have no function other than to rapidly maneuver space assets to new stations. The fleet would have its own dedicated guidance, navigation, and control and telemetry/command system. The tugs themselves could also be refuelable on orbit to extend their lifetimes. Of course some combination of the above approaches for slow or rapid maneuvering and repositioning could be adopted to provide more robust and flexible capabilities, better survivability, and longer life. Recommendation 5-1. DoD should support extensive basic research and technology projects for various in-space propulsion thruster concepts and for in-space electric power generation and energy storage. This fundamental long-range support need not be tied to any specific mission or platform requirement. The current range of technical opportunities is so great that progress will be directly proportional to annual
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs resource allocations over the next 10 years. The committee estimates that at least $20 million per year should be considered as a yearly allocation in these areas. CURRENT STATE OF THE ART IN ON-ORBIT PROPULSION This section of the report discusses the current state of the art in on-orbit propulsion systems. Newer technology in progress at various government and industrial organizations is presented in the next sections. Chemical Propulsion The conventional chemical liquid propellants now in use are either monopropellant or bipropellant. Liquid bipropellant systems are better performers but are more complex and deliver a fuel and oxidizer mixture that reacts chemically in the combustion chamber. Monopropellant systems provide a single propellant that decomposes at the catalyst bed of the combustion chamber. Widely used, highly reliable state-of-the art chemical systems are the monopropellant hydrazine (N2H4) and bipropellant propulsion systems such as mixed oxides of nitrogen (MON) and monomethylhydrazine (MMH). For orbit circularization and station acquisition, bipropellant engines using MON/N2H4 are also in use. Monopropellants Monopropellant hydrazine thrusters have typical performance characteristics as follows: Thrust range: 0.025-125 lbf Isp range: 225-239 lbf-sec/lbm Restart capability: 750,000 starts at 50 lbf Pressure operating range: 350 psia blowdown at 100 psia Radiative thermal control There are three manufacturers of catalytically decomposed monopropellant hydrazine engines in the United States: Aerojet, in Redmond, Washington; American Pacific, in Niagara Falls, New York; and Northrop Grumman Space Division, in Redondo Beach, California. Bipropellants The bipropellant chemical propulsion system MON/MMH has typical performance characteristics as follows: Thrust range: 0.4-5 lbf Isp range: 250-295 lbf-sec/lbm Restart capability: multiple Pressure operating range: 350 psia blowdown at 100 psia Radiative thermal control In this same thrust class, an innovative bipropellant thruster, secondary combustion augmented thruster (SCAT), has been flight qualified and flown. This thruster operates in the bipropellant mode on MON/N2H4 until the oxidizer is expended and then operates as a monopropellant thruster until all the fuel is expended. Northrop Grumman is the sole supplier for this engine. The bipropellant chemical propulsion systems MON/MMH and MON/N2H4, high thrust, are used in liquid apogee engines. They have the following typical performance characteristics:
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs Thrust range: 100-110 lbf Isp range: 305-326 lbf-sec/lbm Restart capability: multiple Engine inlet operating pressure: 250 psia Radiative/film thermal control MON/MMH liquid apogee engines are typically used in combination with the low-thrust MON/MMH thrusters used for on-orbit propulsive functions. MON/N2H4 liquid apogee engines are advantageous for spacecraft propulsion systems that use monopropellant hydrazine or electrothermal hydrazine or hydrazine arcjet thrusters for on-orbit propulsive functions. The same three manufacturers make low- and high-thrust bipropellant engines: Aerojet, in Redmond, Washington; American Pacific, in Niagara Falls, New York; and Northrop Grumman Space Division, in Redondo Beach, California. Electric Propulsion The expanding range of spacecraft sizes and the changes in the commercial spacecraft industry environment have been presenting new challenges to the chemical propulsion community. There has been a clear need for higher performance propellants and/or thrusters. The advent of power-rich spacecraft architectures provides opportunities to take advantage of various propulsion options that can provide both high power and high Isp. Reducing an onboard propulsion system’s wet mass requirement can either decrease total spacecraft mass or increase payload capacity. In addition, greater demands can be placed on the propulsion system, including more calls for repositioning or longer duration orbit maintenance, increasing useful life. Another option enabled by a reduced wet mass might be a stepdown to a lower-weight-class launch vehicle. These performance enhancements, which are of great interest to commercial satellite owners, are also desirable for military satellites. The propulsion industry has accepted these challenges and is transitioning to electric propulsion. The extent to which the commercial satellite industry has embraced electric propulsion is evident from Figure 5-2, which shows all the operational satellites using electric propulsion as of June 2006. The thrusters shown in the figure, which will be discussed in more detail in the following paragraphs, are of several types: electrothermal hydrazine thrusters (EHTs), hydrazine arcjets, gridded ion thrusters, Hall-effect thrusters (HETs), and pulsed plasma thrusters (PPT).
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-2 Operational satellites with electric propulsion. SOURCE: Aerojet (2006). Electrothermal Thrusters Starting with the implementation in the 1980s of EHTs on INTELSAT V and RCA AstroElectronics communication satellites, a 30 percent improvement in performance, from an Isp of 225 sec to an Isp of 295 sec, was achieved with this thruster type, which electrically heats the decomposition products of catalytically decomposed monopropellant hydrazine to higher chamber temperatures. Without the complexity of carrying an oxidizer on board, this Isp is competitive with low-thrust bipropellant systems (0.4 to 5 lbf), which provide an Isp of 295 sec. TRW Space and Communications (now Northrop Grumman), Redondo Beach, built the EHTs for INTELSAT V. At present it does not manufacture this thruster type. The EHTs that are currently in use (see Figure 5-2) are manufactured by Aerojet Redmond. A schematic of the Aerojet Redmond EHT and its characteristics are shown in Figure 5-3.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-3 Aerojet MR-501B electrothermal hydrazine thruster (EHT). SOURCE: Aerojet (2004b). Arcjets In the early 1990s, Lockheed Martin utilized a new thruster, the hydrazine arcjet for North-South station keeping (NSSK) on its geostationary orbit satellites. The early R&D (through preflight qualification) on the hydrazine arcjet was done at NASA Glenn Research Center. The current production arcjet thrusters are manufactured by Aerojet Redmond. The Lockheed Martin series 7000 satellites use the Aerojet MR 509 hydrazine arcjet system (1.8-kW power level, Isp of 502 sec). The arcjet continues to evolve with the latest Lockheed satellite bus, the A2100 satellites, which utilize the MR-510 arcjet system (2.2-kW, 582-sec nominal Isp thrusters) for NSSK. Again, the arcjet thruster takes advantage of the higher satellite power available to substantially increase the performance over catalytic hydrazine (Isp = 225 sec to Isp = 570 to 600 sec). A schematic of the MR-510 and its characteristics are shown in Figure 5-4.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-4 MR-510 arcjet thruster and cable assembly. SOURCE: Aerojet (2004b). Ion Thruster Systems In the late 1990s, Boeing Electrodynamics (formerly Hughes Space and Communications) successfully introduced the Xenon Ion Propulsion System (XIPS), a xenon propellant gridded ion thruster, on its BSS 601HP and BSS 702 commercial communications geosynchronous G satellites. The XIPS thruster schematic is shown in Figure 5-5. Boeing Electrodynamics XIPS thruster technology was recently purchased by L-3 Communications’ Electron Technologies, Inc., which presently manufactures thrusters for both satellites. FIGURE 5-5 XIPS thruster schematic illustration. SOURCE: Encyclopedia of Astrobiology, Astronomy, and Spaceflight (Undated). The thruster consists of a discharge hollow cathode, three-ring magnetic cusp confinement, a three-grid accelerator, and neutralizer hollow cathode. The three-grid accelerator used in the 25-cm thruster utilizes shaped molybdenum grids with approximately 11,000 apertures to produce the high perveance (72 pervs at full power) xenon ion beam. The XIPS 25-cm ion thrusters and the associated power supplies operate in two modes: 2.2 kW for typical on-orbit functions and 4.4 kW for raising the orbit. The high-power mode utilizes about 4.5 kW of bus power to produce a 1.2-kV, 3-Å ion beam. The thruster in this
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs mode produces 165 mN thrust at an Isp of about 3,500 seconds. The high-power mode is used exclusively for the orbit insertion phase, which greatly reduces the amount of chemical propellant carried by the spacecraft for this task. Nearly continuous operation in the high-power mode for 500 to 1,000 hours is required, depending on the launch vehicle and satellite weight. A low-power mode, in which the thruster consumes about 2.2 kW of bus power, is used for the station-keeping function. In the low-power mode, the beam acceleration voltage is kept the same, and the discharge current and gas flow are reduced to generate a 1.2-kV, 1.43-Å beam. In this mode, the thruster produces 79 mN of thrust. Since the beam voltage remains unchanged for the high-power mode and the thruster mass utilization efficiency is nearly the same, the specific impulse degrades only slightly compared to the high-power mode, to about 3,400 sec. Typical performance parameters of the 25-cm thruster used on the BSS 702 satellites are summarized in Table 5-1. TABLE 5-1 Typical Parameters of the 25-cm XIPS Thruster Low-Power Station Keeping High-Power Orbit Raising Active grid diameter (cm) 25 25 Average Isp (sec) 3,400 3,500 Thrust (mN) 79 165 Total power consumed (kW) 2.2 4.5 Mass utilization efficiency (%) 80 82 Typical electrical efficiency (%) 87 87 SOURCE: Goebel et al. (2002). The state of the art on the XIPS is described by Chien et al. (2006). The Boeing 702 spacecraft has a chemical propulsion liquid apogee engine, but use of the high-power mode of XIPS in the orbit insertion phase greatly reduces the wet mass carried by the spacecraft for this task. The high Isp of the ion engines for NSSK results in an additional large saving in propulsion system wet mass over on-orbit systems, which use mono- or bipropellants for this function. The military Gapfiller satellite, which launches in 2007, will use the 25-cm thruster version of XIPS. Aerojet also has a major development effort in xenon ion thruster system technology. It is completing the thruster, propellant management, and digital control systems designs on the 0.5-7.5 kW NASA Evolutionary Xenon Thruster (NEXT) effort, led by NASA Glenn Research Center. L-3 Communications Electron Technologies, Inc., is developing the power processor. The NEXT system, when qualified, will provide much greater capability for Discovery-class solar electric propulsion missions. The 40-cm NEXT will also be available for other spacecraft applications. Hall-Effect Thrusters In 1990, the Science and Technology Directorate of the Ballistic Missile Defense Organization (BMDO) took the lead in identifying advanced spacecraft propulsion technology developed in the former Soviet Union with potential applications for U.S. government and commercial missions. It identified the Russian Hall thruster technology as being particularly promising (Sankovic et al., 1997). In 1971, the Russians flew the first Hall thruster—sometimes called a stationary plasma thruster (SPT)—on the METEOR spacecraft. Over the next two decades several dozen 0.66-kW SPT-70 thrusters were used operationally in space. BMDO procured three versions of the 1.5-kW Hall thruster for evaluation: (1) SPT 100, (2) T100, and (3) TAL D-55. This procurement provided three sources of Hall thrusters for U.S. propulsion companies. Hall system thruster development has gone forward at Aerojet Redmond, Busek, Loral Space Systems, and Pratt & Whitney. Use of Hall thrusters for satellite NSSK promises great savings in wet mass over mono- or bipropellant chemical propulsion systems. An overview of the underlying physics is available in Kaufman (1985).
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs A typical propellant for a Hall thruster is a high-molecular-weight inert gas such as xenon. A power processor is used to generate an electrical discharge between a cathode and an annular anode, through which the majority of propellant is injected. A critical element of the device is the incorporation of a radial magnetic field, which serves to impart an azimuthal drift to the electrons coming from the cathode and to retard their flow to the anode. The azimuthally drifting electrons collide with the neutral xenon, ionizing it. The xenon ions are then accelerated electostatically from the discharge chamber by the electric potential maintained across the electrodes by the power processor. The velocity of the exiting ions, and hence the Isp, is governed by the voltage applied to the discharge power supply and is typically 15,000-16,000 m/sec at 300 V. The first flight of a Hall thruster on a U.S. spacecraft was in 1998 on Space Technology Experiment (STEX), a Naval Research Laboratory spacecraft. In 2004, Loral launched the Mobile Broadcasting Satellite (MBSAT), a geosynchronous satellite that uses four Faekel SPT-100 Hall thrusters for NSSK. The performance characteristics of the SPT-100 class of thrusters are shown in Table 5-2. TABLE 5-2 Characteristics of SPT-100 Hall-Effect Thrusters Characteristic Value Propellant Xenon Thrust (mN) 80 Power (kw) 1.35 Isp (sec) 1,600 Efficiency (%) 50 Life (hr) >7,000 SOURCE: Sankovic et al. (1993). Aerojet Redmond has completed flight qualification of the BPT-4000, shown in Figure 5-6, for the Lockheed Martin build of the Air Force advanced EHF satellites. These Hall thrusters will be operated at two thrust levels, a high thrust for partial orbit transfer and lower thrust for station-keeping requirements. Launch date is projected as the fourth quarter of 2006. FIGURE 5-6 Aerojet dual mode BPT-4000 Hall effect. SOURCE: Aerojet (2005b). The availability of flight-qualified, flight-proven ion and Hall thrusters can be expected to increase the manifesting of these technologies because they have higher Isp than chemical propulsion systems. Each type of electric-powered thruster has its area of applicability. Ion engines can deliver higher Isp and are
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs well suited to missions with high ΔV requirements. In addition to satellite NSSK and partial orbit transfer requirements, Hall thrusters would be suitable for applications such as Earth transfer missions. In fact, the European Space Agency’s (ESA’s) Small Missions for Advanced Research in Technology 1 (SMART 1) mission has already successfully implemented this technology. SMART 1 is a small lunar orbiter that was launched in September 2003 as an auxiliary payload on the Ariane 5. It uses a 1.4-kw Hall thruster and reached the first moon orbit in December 2004. Due to the mass limitation of the spacecraft and the consequent limitation in the electrical power, the thruster used on SMART 1 is a scaled-down version of the PPS-1350 thruster developed and qualified by SNMECA (France) for geosynchronous missions. SMART 1 used its thrusters in a variable power mode (450 to 1,220 W) in this application, which serves as a benchmark for other Earth-to-orbit (ETO) missions using electric propulsion. Two companies, Moog and Vacco Industries, are leading efforts to produce propellant management components and feed systems for flight electric propulsion systems. They are also developing next-generation designs that will trim the weight of the propellant management system. Moog was the supplier of the xenon propellant management assembly that is flight operational on the Loral MBSAT. Currently, Vacco Industries provides the propellant management system components for the BSS 702 satellite XIPS. It is in the process of qualifying highly integrated chemically etched propellant management (CHEMS) xenon-latch-valve modules for the Lockheed Martin advanced EHF satellite. To ensure broader application of Hall thrusters, and ion thrusters as well, more attention needs to be paid to developing the components of the entire electric propulsion subsystem, which includes not only the thruster but also the propellant feed system and the power processing unit (PPU). Historically the PPU has been the dominant cost driver for electric propulsion systems because of the requirement for heavy power converters and thermal management systems. Aerojet Redmond designs and builds high-power converters to support the electric propulsion subsystems it manufactures. It is also working on the development of solar-electric direct drive, i.e., using a high-voltage solar array to provide power directly to a Hall thruster at voltages needed to drive thruster discharge. Qualification of the solar-electric direct drive would greatly reduce the cost and weight of a Hall electric propulsion system, while reducing array size. Reduction in array size results in an added savings in spacecraft weight. The potential payoff for direct drive makes this a goal extremely worthy of pursuit. Additional weight savings could be obtained with direct drive using power from advanced solar arrays now available commercially, such as ENTECH/ABLE’s Solar Concentrator Arrays with Refractive Linear Element Technology. Such an array provided the 2.7-kW power source for the successful Deep Space 1 basic mission and its extended mission to the comet Borrelly in 2001 (Jones et al., 1996). ENTECH/ABLE also has available a next-generation array, which is a stretched-lens array. The synergy of coupling flight-proven, advanced-array technologies with direct drive for Hall thrusters needs to be explored. To satisfy increasing demand, a larger industrial base is needed than now exists for the manufacture of electric propulsion systems and components. L-3 Communications Electron Technologies, Inc., and Aerojet Redmond appear to be the only commercial sources for gridded ion thrusters and .power converters. Aerojet is the only U.S. source that has qualified Hall thruster hardware. In addition to the dual-thrust 4.5-kW Hall thruster under contract to Lockheed Martin for the EHF satellite, Aerojet has fabricated and tested a 2.2-kW Hall thruster flight prototype unit. Busek is developing low-power Hall thrusters with IHPRPT funding; Busek designs are discussed under “Electric Propulsion” in the IHPRPT Targets section. Excellent R&D on ion and Hall thrusters and power-conditioning units is being conducted at NASA Glenn Research Center, but there needs to be more transfer of technology to companies that can build the product. Propulsion for Microsatellites There is presently a substantial interest in microsatellites: satellites with masses from 50 to 100 kg. For example, microsatellites could be used for missions requiring formation flying, precise attitude control, or trajectory correction. Design and implementation of the micronewton-thrust propulsion systems required for precision control is especially challenging given the mass, volume, and power
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs constraints that come from the satellite’s small size. The most promising technologies under investigation to date for near- and intermediate-term military applications appear to be micronewton PPTs, colloid thrusters and ion thrusters, and both ion grid thrusters and Hall effect thrusters. The following discussion describes the characteristics of a colloid thruster and several PPTs that have been flown or are nearly flight ready. The investigation of micropropulsion is ongoing at NASA’s Jet Propulsion Laboratory (3-cm-diameter xenon ion thruster), at L-3 Communications Electron Technologies (8-cm-diameter xenon ion thruster), at Stanford University (micro-Hall thruster), and at many other university sites, but these microelectric propulsion concepts are not sufficiently far along to be considered for near- or intermediate-term flight application. Pulsed Plasma Thrusters The Aerojet PPT, designated PRS-1, was successfully flown on the NASA Goddard Earth Observing EO-1 spacecraft. This PPT relies on the Lorentz force generated by an arc passing from anode to cathode and the self-induced magnetic fields to accelerate a small quantity of chlorofluorocarbon propellant. Teflon has been used as the propellant to date. Pulsed electromagnetic thruster systems consist of accelerating electrodes, an energy storage unit, a power conditioning unit, an igniter supply, and a propellant feed system. During operation, an energy storage capacitor is first charged to between 1 and 2 kV and an ignition supply is then activated to generate low-density plasma, which permits the energy storage capacitor to discharge across the face of the Teflon propellant bar. The peak current level is typically between 2 and 15 kA, and the arc duration is between 5 and 20 microsec. The pulse cycle can be repeated at a rate compatible with the available spacecraft power. The propellant feed system consists of a negator spring that pushes the solid Teflon bar against a stop on the anode electrode. The characteristics of the PRS-1 pulsed plasma system, flown on the EO-1 spacecraft, are shown in Table 5-3, along with those of a microthrust PPT system built by the University of Washington.1 TABLE 5-3 PPT Performance Characteristics Characteristic EO-1a Dawgstarb Maximum input power (W) 70 (1 thruster; EO-1 operations); 100 design 15.6 (2 thrusters; slow charge) 36.0 (2 thrusters; fast charge) Thrusters per system 2 8 Total system impulse (N-sec) 1,850 (EO-1 propellant load) >15,000 (system life) >1,500 Impulse bit (μN-sec) 90-860, throttleable 66 ± 4 Pulse energy (J) 8.5-56, throttleable 4.9 Maximum thrust (μN) 860 (EO-1); 1.2 (design) 200 (high-speed mode) Specific impulse (sec) 650-1,350 332 ± 40 Thrust to input power ratio (μN/W) 12.3 9.7 Total mass (kg) 4.9 (2 PPTs, 1 PPU, and propellant) 4.2 (8 PPTs, 1 PPU, and propellant) Propellant PTFE PTFE Propellant mass (kg/thruster) 0.07 (as fueled) 0.07; 0.56 kg/system aSOURCE: Benson et al. (1999). bSOURCE: Rayburn et al. (2000). The PRS-1 has demonstrated control of the spacecraft pitch with the momentum wheels completely disabled, including during image acquisition with the Advanced Land Imager instrument. On-orbit tests have demonstrated no detectable electromagnetic interference with the spacecraft, spacecraft communications, or the payload instrument. 1 A prototype of this system was to be flown on the Dawgstar nanosatellite, which is currently in storage waiting for a launch opportunity.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs Storable liquid propellant systems provide total thrust flexibility at high combustion efficiency. Liquid propellants for tactical missiles have not found favor with any of the Services and will not be discussed further in this report. Gelled propellant motors provide total thrust flexibility at high combustion efficiency and meet IM requirements by storing fuel and oxidizer in separate tanks. The Army’s gelled propellant controllable thrust systems will be discussed below. Solid propellant motors using a variable area nozzle can provide some flexible thrust capability and sensitivity of the solids at lower operational temperature. A minimum-signature, Class 1.3 propellant with a variable area nozzle demonstrated a 50-sec test with multiple thrust variation, no gas leakage, and minimal hardware degradation. Trade studies have shown that pulsed solids can increase range 15 to 30 percent over boost-sustain systems. Static testing has shown successful pulse motor tests without barriers and low-pressure smoldering with a return to full-boost thrust. In addition to the Army in-house program at Huntsville, Aerojet and ATK have contributed to recent advances in controllable thrust solid technology. Controllable thrust systems are being considered for upgrades for three emerging missile systems in the FY08-FY10 time frame. Insensitive Munitions. The challenge for solid propellant formulators is that high-energy propellants generally yield more violent IM test results. There is no such thing as an IM propellant formulation. IM, however, is a system issue and can be achieved by case material and engineering design. Current minimum-signature solid propulsion development includes the development of less-sensitive propellant formulations, case materials, and case venting. New Materials. As a part of the Ordinance Environmental Program, the Army is developing a non-Pb, minimum-signature, Class 1.3 solid formulation. The goals for the new solid propellant formulation include a specific impulse range of 230-240 lbf-sec/lbm, a burning rate of 0.2-0.6 in./sec at 1,000 psi, a burning rate exponent >0.5, and a dependence of motor chamber pressure on initial propellant temperature of >0.15 percent per Fahrenheit degree. The plan is to reach a TRL of 6 by the end of FY08. Hybrid Motors U.S. Army Hybrid Controllable Thrust Motor A hybrid propulsion system is being evaluated as an IM-compliant alternative to nonammonium perchlorate solid propulsion systems. As shown in Figure 5-9, two types of hybrid rockets have been considered: (1) a classical hybrid rocket (conventional design) in which liquid or gelled oxidizer is injected into the port(s) of the solid-fuel grain or the fuel-rich propellant grain for combustion and (2) a gas-generator type of hybrid rocket in which the fuel-rich solid propellant grain burns in its own combustor and the discharged products are further burned with the oxidizer-rich gases in a post combustor. In some special cases, an inverse hybrid can be considered, in which the solid grain is made of oxidizer-rich material and the injected liquid is a fuel-rich material. This design option is rarely used. The separation of oxidizer and fuel reduces the sensitivity of the propulsion system to external stimuli for improved IM characteristics developed under internal independent research and development (IR&D) funding and increases the TRLs of numerous hybrid-based systems.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-9 Hybrid missile concepts. SOURCE: Morrison (2005). Lockheed Martin Space Systems Company Lockheed Martin Space Systems Company (LMSSC) has worked on hybrid propulsion technologies since 1989. Its initial studies were focused on replacing the solid rocket boosters on the space shuttle after the Challenger disaster. It worked with American Rocket Company (AMROC) during the DM-01, DM-02, and hybrid technology options project (HyTOP) motor developments, which eventually led to the hybrid propulsion development program (HPDP). Within the HPDP, LMSSC tested numerous technologies that were hybrid-based systems. Hybrid Technology Performance The Isp of the hybrid propellant combination used for the LMSSC FALCON stages at various expansion ratios is shown in Figure 5-10. Because the fuel is inert, launch vehicles or missiles that use these propellant combinations can achieve good performance and gain the benefits of having a nonexplosive propellant combination. LMSSC claims that its staged combustion system helps this propulsion system to achieve the maximum efficiency possible with hybrids. Data from testing indicate that the system is efficient enough and stable enough to be competitive with liquid- and solid-based propulsion systems.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-10 Hybrid motor performance at different expansion ratios. SOURCE: Lockheed Martin Space Systems Company. Current Hybrid Technology Lessons learned from data gathered for AMROC, HyTOP, IR&D and HPDP indicated that heat had to be added to the forward end of a hybrid motor to ensure stability and high efficiency in hybrid motors. LMSSC developed and patented an active approach, the staged combustion system (U.S. Patent 5,794,435), to accomplish this task. Figure 5-11 shows the staged combustion concept for hybrid motors. FIGURE 5-11 Staged combustion approach for hybrid propulsion systems. SOURCE: Lockheed Martin Space Systems Company. To demonstrate the effectiveness of this concept, a series of tests were performed at various thrust levels, ranging from 1,500 lbf to 250,000 lbf. The first test of the series served as an unstable baseline test that was ignited using triethylaluminum/triethylborane (TEA/TEB). The same motor was retrofitted with
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs the staged combustion system and showed significantly higher performance and stability within 2.5 percent of the average chamber pressure (typical solid propulsion stability is 5 percent of the average chamber pressure). Figure 5-12 shows the results from the testing of these 1,500 lbf motors at MSFC during the HPDP. FIGURE 5-12 Staged combustion approach provided stable combustion and increased performance in the 11-in. diameter hybrid motors. SOURCE: Lockheed Martin Space Systems Company. A unique approach, which is currently used on LMSSC’s FALCON SLV first- and second-stage hybrid motors, is to employ a multirow, multiport hybrid fuel grain that allows for increasing the diameter of the vehicles beyond what could have been achieved with the single-row technology used during the HPDP. An enabling technology for this type of design is an improved hybrid fuel, which was developed under IR&D by Lockheed Martin, allows for approximately 10 times more propulsive power than the hybrid sounding rocket fuel and enables the fuel webs to become very thin prior to failure during the burn, which, in turn, allows for minimum residuals. LMSSC conducted several successful tests that advanced the state of the art for hybrid rocket motors. The most significant of these tests was a 120-sec test firing of a 30,000-lb-thrust hybrid motor at AFRL on June 10, 2005. The test, conducted as part of the FALCON SLV program, was the second SLV hybrid motor firing that LMSSC carried out that year at AFRL test stand 2-A. The company conducted six more tests in 2006. These tests demonstrated the use of a hybrid rocket fuel-rich motor in a gas generator to drive a pump turbine and tested new high-performance fuels, advanced nozzle materials, and the dramatic improvements in performance and stability enabled by the GOx-rich upstream staged-combustion preburner. Enabling Technologies for Hybrid Propulsion Systems Enabling technologies for (1) storable oxidizers; (2) minimizing residual fuel grain; (3) motor insulation compatible with hybrid combustion products; and (4) materials for the nozzle throat that inhibit erosion will be important for hybrid propulsion systems. For example, the hybrid motor technology planned at Lockheed Martin includes the development and use of storable oxidizers, higher-density-fuel formulations, programmable start/stop and restart capability, and mission-specific ability to throttle. Filled ethylene propylene diene monomer (EPDM) insulators have been proved adequate for solid propulsion systems. With hybrid systems, rubber-based insulators act as fuel and have been shown to erode relatively fast when exposed to hybrid exhaust constituents. Insulation materials compatible with hybrid combustion products need to be improved to accomplish the run-to-empty goal. Future hybrid motor insulators will need to serve as a structural element during the initial burn, when the chamber pressure loads are highest, and will need to withstand erosion when exposed. Material testing in a relevant environment will enable minimum fuel residuals and lower the overall mass of inert insulation material. Test data indicate that the nozzle throat materials typically used for solid propulsion systems, such as three-dimensional carbon-carbon and ATJ graphite, erode fairly quickly in a high-pressure environment
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs containing the combustion products of hybrid motors. Rapid nozzle throat erosion 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. Nozzle materials that are more compatible with hybrid propulsion systems will have to be identified and developed. Alternatively, cooling techniques, such as film cooling with the fuel or the on-board oxidizer, could be employed to reduce throat erosion rates well below 5 mil/sec. This will be necessary for long-duration missile motor burns. Residual hybrid fuel on an upper stage translates directly to payload mass, which contributes to the physical size of a hybrid missile propulsion system as compared with a solid or liquid stage. Reducing the amount of fuel residual will decrease the physical size of hybrid propulsion systems, making them comparable to other systems. Future hybrid fuel grains will need to be tailored for the fuel webs to merge as a function of time to drive the fuel residual share well below 2 percent of the total fuel on board. Also, fuel structural strength will need to be improved and long-duration testing will be required to advance the current state of the art. Gelled Propellant Motors Gelled propellant motors can provide total missile thrust flexibility at high combustion efficiency and meet IM requirements by storing fuel and oxidizer in separate tanks. Controllable thrust can provide significant system benefits. It can provide extended range and shorter time-to-target at midranges in a single system and can reserve propellant energy for endgame performance. Gelled propellants can meet operational and handling requirements. Gel propulsion for airborne missiles has the potential to be inherently insensitive to IM threats because the fuel and oxidizer are stored in separate tanks. Gel propulsion systems have passed bullet impact, slow cook-off, fast cook-off, and shaped-charge jet IM tests. Logistics costs can be reduced by substituting a single controllable-thrust missile for currently deployed single-use systems. They are flexible enough to meet evolving airborne missile system requirements. U.S. Army, Huntsville, Alabama Gel propulsion systems have flown successfully during two tests of systems to integrate future missile technology and have demonstrated operability at minus 40°C while maintaining 96 percent of ambient thrust and 97 percent of its ambient-density specific impulse. A throttling gel engine using a passive pintle demonstrated a turndown ratio of 12:1 while maintaining >98 percent Isp efficiency. In addition to the Army in-house program, Northrop Grumman, Aerojet, Stone Engineering, and the ORBITEC-CFDRC team have contributed to recent advances in gel propulsion technology. Northrop Grumman Corporation Solid propellant motors have been almost universally used in tactical aircraft-borne missiles. Generally these solid motors are simple fixed-thrust booster stages. Flexible flight profiles and endgame maneuvers are not well suited to solid propellant motors. Although solid propellant motors with variable plug nozzles can be throttled and can sometimes be shut down and restarted, their operating profile throughout a long fly-out mission is quite limited. Storable liquid propellant rocket motors have much more flexibility in operating profiles. However, handling and leakage have always been of concern. Also, liquid propellant systems tend to have lower density. Because aircraft-borne missiles are usually volume constrained by aircraft configuration, lower density can mean the aircraft is not able to carry as much total impulse. Although the Isp of liquid propellant can be higher than that of the best solid propellants, their density impulse may only just match that of solids. The development of gelled-liquid propellants over the last 20 years has provided attractive options for very demanding flight profiles. Gel-propellant formulations with very-high-density specific impulse have been demonstrated. These gels have the physical consistency of heavy toothpaste and can be stored
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs indefinitely with no leakage. In fact, they can be frozen at extremely low temperatures and then rapidly thawed out for use with no detrimental effects. They also do not have the time-dependent deterioration issues experienced by most solid propellant grains and case bonds. Under high shear loads, gelled propellants behave like normal liquids. They can be throttled, shut off, and restarted when using the right type of injectors. They can be pulsed with variable off times up to almost any value demanded. Design criteria for gelled-propellant rocket systems for missiles of almost any size have been validated at Northrop Grumman Corporation. The critical components of a system are these: A special injector that permits throttling and no-dribble shutoff and restart, A coaxial-piston bipropellant tank system, and A gas-generator-piston pressure system. The key technology for enabling the flexible multistart and thrust profile tailoring is a face shutoff injector. For this application, a single central-element pintle injector turns out to be ideal. With a single sleeve, both the oxidizer and fuel can be throttled while maintaining absolute control of the mixture ratio at all thrust levels, and the engine can be shut off at the face so that neither gelled propellant can evaporate at shutdown. This injector is essentially the same as the type used for the Lunar Module descent engine. It delivers efficient combustion with absolute dynamic stability. A cutaway illustration of an injector with a face shutoff configuration is shown in Figure 5-13. Finding 5-4. The very high packaging density and optimized thrust profiles possible with these gelled propellant systems can significantly increase the range for a given fly-out missile envelope. This can also be an advantage for in-space fly-out missiles or responsive space tugs. FIGURE 5-13 Controlling sleeve shown in shutoff position (center flow is oxidizer; flow around the sleeve is fuel). SOURCE: Northrop Grumman Corporation. Recommendation 5-4. DoD should ensure that the development of advanced tactical missiles, responsive global-reach missiles, and antiballistic missiles (ABMs) satisfies four key requirements: effective energy/trajectory management; higher-energy-density performance; minimum smoke exhaust; and insensitive propellants. The S&T part of the DoD/Air Force strategic plan for missiles should focus on the technologies and design criteria necessary to meet these goals. The committee’s estimate of annual
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs funding that would be required to make reasonable progress in establishing advanced capabilities in these areas is $20 million to $30 million. OPPORTUNITIES FOR TRANFORMATION IN ACCOMPLISHING RESPONSIVE GLOBAL REACH AND ABM MISSIONS There are a couple of opportunities for transforming the means by which certain responsive global reach and ABM missions can be achieved. Two concepts for space access launch vehicles were described and discussed in Chapter 4. The BAE Systems concept utilizes self-contained air-based vertical launch (ABVL) modules. The second system, under study by DARPA and NASA, makes use of a multimission modular vehicle (MMMV). Both air-based launch platforms could transport rocket-powered missiles to high-altitude launch points at desired geographic locations. Both would enable tremendous flexibilities in launch time and azimuth (orbital inclination) for missiles for ABM missions, tactical support, or long-range global strike. Such platforms could provide a faster response to emerging threats than is available today. AIR-BASED VERTICAL LAUNCH CONCEPT Launching missiles from a flying aircraft platform could dramatically improve the delivery of a warhead and decrease the time to target. High-altitude air launch allows a rocket to bypass the initial parts of a ground-launch trajectory. This part of the trajectory is where the missile’s orientation is mostly vertical and where the gravity × time losses in delivered impulse are the greatest. These losses are compounded by the reductions in thrust and specific impulse that rocket engines experience at high ambient pressures. In addition, the integrated drag losses between the ground and 40,000 feet are great. The combined effects could reduce velocity by about 3,500 feet per second. Also, even if a ground launch takes place in the best geographic location, which is not likely, the time to reach 40,000 feet can be 20 to 40 seconds. This time difference could be crucial for a boost-phase ABM mission. Furthermore, vertical launch vs. air drop can cut many seconds off the fly-out time to target from a platform at a given location, velocity, and altitude. In the BAE concept, missiles are preloaded in a self-contained, installable vertical launch module. The module can accommodate many missiles of several types. This could permit multiple targeting for strike and tactical battlefield support. Some potential mission capabilities of a joint strike air network architecture are depicted in Figure 5-14.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs FIGURE 5-14 Multimission global shield using some precision guided missiles. SOURCE: Smith (2006). Finding 5-5. Much of the launch dynamics and environment of an air-based vertical launch is very different from a ground-based launch, a Pegasus launch, or the candidate FALCON AirLaunch vehicle, described in Chapter 4. Characteristics for candidate missile propulsion systems (including parallel booster or strap-on combinations), along with propulsion technologies such as propellant combinations (solids, storable liquids, gelled combinations, and 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 ABVL for global strike and near-space military applications. To exploit this launch concept, one of first technology demonstration efforts should obtain some data on launch environment dynamics needed to carry out system trades. Recommendation 5-5. The Air Force should sponsor basic missile/environment dynamics measurements and detailed system engineering studies to fully understand the transformational potential of utilizing air-based vertical launch concepts for various types and sizes of prompt-response military missiles. The propulsion technologies that need to be evolved to take full advantage of such launch platforms should be identified and developed.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs MultiMission Modular Vehicle Concept The new multipurpose airframe design under consideration by DARPA for future applications, currently designated the multimission modular vehicle (MMMV), is described in Chapter 4. The airframe is designed in such a way that the centerline payload could be either a self-contained launching pod for multiple medium-sized missiles or a single large missile. Either configuration could also be equipped with folded rotor blades for emergency separation or self-transport. The MMMV concept could also provide a transformational missile launching capability for large or small missiles. The aircraft can be configured with a specialized missile pod. Like the airborne vertical launcher it could transport larger rocket-powered missiles to high-altitude launch points at optimum geographic locations. The versatility of a removable centerline, self-contained, operational launcher module is key to this system’s high load capacity and operational flexibility (see Figure 5-15). Finding 5-6. Missile configurations and propulsion technologies would need to be optimized to take full advantage of the transformational potential of aircraft configured to launch missiles at high altitudes. Some of the propulsion technologies 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, T/W profiles, thrust vector control, and rocket plume impingement patterns. FIGURE 5-15 MMMV with missile launcher. SOURCE: NASA MFSC. Recommendation 5-6. The Air Force and DoD should sponsor a detailed system engineering study of using the Multi-Mission Modular Vehicle air-based launch system for medium-sized vehicles could be combined in combination with a study on using air-based vertical launch for small vehicles, ensuring they are focused on Air Force/DoD mission success optimization criteria. The studies would identify the propulsion technologies (modifications or new concepts) that should be evolved in order to take full advantage of such air-based launch platforms for operationally responsive missions. Critical Enabling Technologies Critical enabling technologies for substantial improvement of missile propulsion operational capabilities should be redoubled to find new energetic propellants and heat- and chemicals-resistant
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs materials that have the potential to more fully enable DoD ground and airborne missile applications. These two areas are discussed below. New Energetic Propellants Prospects for the search for energetic yet insensitive propellants in the near term seem poor. Monopropellants with higher density Isp may evolve first, but even if one is validated it can be expected to take many years to establish a reliable industrial product capability at an acceptable cost. Chamber and Nozzle Throat Materials A major problem limiting the future use of any new energetic propellants even if they become available is the lack of materials that are resistant to chemical attack and to erosion at high temperatures. The high temperatures achieved by energetic propellants will produce the same molecules as are produced by other fuels, including CO2, H2O, N2, and CO. The requirement for low erosion materials is a result of the higher temperatures achieved by these propellants. Finding 5-7. If the DoD and the Air Force are going to realize any transforming options in the specific performance profiles of tactical missiles in the far-term, a well-funded, continuous effort in energetic fuels and resistant materials is required. Recommendation 5-7. DoD and the Air Force should fund the search for new high-energy propellants and development of very-high-temperature, chemical-attack-resistant, low-erosion-rate materials. FINAL OBSERVATION Actual funding levels for technology programs such as IHPRPT and for sustaining and improving the engineering on tactical and strategic missiles have dropped well below the original planned funding levels. This limits the accomplishments of propulsion improvement efforts and minimizes the ability to train the next generation of designers and production specialists. Personnel demographics predict the retirement of individuals with critical skills in the development and production of large missiles and launch vehicles in this same time frame. The consequences of this situation have been eroding U.S. aerospace capability for many years. Unless a serious commitment to reversing this trend is made, the ability of the industry to provide the high-quality engineering and production capability necessary to realize the Air Force’s medium- and far-term goals for access to space, in-space operations, and missiles must be considered at risk. REFERENCES Published Benson, S. W., L.A. Arrington, W.A. Hoskins, and N.J. Meckel. 1999. Development of a PPT for the EO-1 Spacecraft. AIAA-99-2276. June. Chien, Kuei-Ru, William G. Tighe, Thomas A. Bond, and Rafawl Spears. 2006. An Overview of Electric Propulsion at L-3 Communications Electron Technologies, Inc. AIAA Paper 2006-4322. Encyclopedia of Astrobiology, Astronomy, and Spaceflight (undated). Available at http://www.daviddarling.info/encyclopedia/X/XIPS.html. Last accessed on June 16, 2006. James, Larry. 2005. HLV study and analysis industry day welcome and introductions,” Presentation to ARES Industry Day, El Segundo, Calif., March 7. Jones, F., D. Murphy, D. Allen, L. Caveny, and M. Piszczor. 1996. SCARLET: High-Payoff, Near Term Concentrator Solar Array. AIAA Paper 96-1021. Kaufman, H.R. 1985. Technology of closed-drift thrusters. AIAA Journal 23(1): 78-87.
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A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs Randolph, Thomas M., John K. Ziemer, Vlad Hruby, Doug Spence, Nate Demmons, Tom Roy, Bill Connolly, Eric Ehrbar, Jurg Zwahlen, Roy Martin, and Chas Gasdaska. 2006. Microthrust Propulsion for the Space Technology 7 (ST7) Technology Demonstration Mission. AIAA Paper 2006-4320. Rayburn, C., M. Campbell, A. Hoskins, and J. Cassady. 2000. Development of a Micro-PPT for the Dawgstar Nanosatellite. AIAA Joint Propulsion Conference. July. Sankovic, J., L.H. Caveny, and P. Lynn. 1997. The BMDO Russian Hall Electric Thruster Technology (Rhett) Program: From Laboratory to Orbit. AIAA Paper 97-2917. Unpublished Aerojet. 2004. Redmond Operations Datasheets. Aerojet. 2006. Redmond Operations Capabilities 2006-H-3029. July. Mike Huggins, Air Force Research Laboratory. “IHPRPT overview,” Background information provided to the committee on April 14, 2005. Michael Morrison, United States Army. “Propulsion overview and roadmap,” Presentation to the committee on June 23, 2005. Scott Smith, BAE Systems. “Air-based vertical launch,” Background information provided to the committee on January 6, 2006.
Representative terms from entire chapter: