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Summer 2011

Aerospike for Increased Space Launch Capability

Carl R. Hartsfield Air Force Institute of Technology

Richard D. Branam

Joshua N. Hall

Joseph R. Simmons III

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Recommended Citation Hartsfield, C. R., Branam, R. D., Hall, J., & Simmons, J. R. (2011). Aerospike Rockets for Increased Space Launch Capability. Air and Space Power Journal, 25(2), 65–73.

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Top Aerospike Rockets for Increased Space Launch Capability Lt Col Carl Hartsfield, PhD, USAF Lt Col Richard D. Branam, PhD, USAF Capt Joshua Hall, USAF Mr. Joseph Simmons*

he US Department of Defense (DOD) loads into the desired orbits, permitting increasingly depends on space assets new missions and capabilities. To overcome for everyday operations. Precision these limitations, the Air Force Institute of Tnavigation; communications; and intelli- Technology (AFIT) conducts ongoing re- gence, surveillance, and reconnaissance sat- search into propulsion technologies ellites are highly leveraged space assets. to improve space launch performance. The launch vehicles that place these satel- Two significant problems hinder space lites in orbit are a major limitation of cur- launch today: launch performance and cost. rent space systems. If higher-performing Performance involves the payload mass that launch vehicles were available, many satel- a vehicle can place into a given orbit, whether lites could accommodate additional capa- low Earth orbit (LEO) or geosynchronous bilities, whether in terms of more sensor Earth orbit (GEO). The Delta IV Heavy, ca- channels, types of payloads, electrical pable of delivering 50,655 pounds into LEO power, or for - or 14,491 pounds into GEO, represents the ing and station keeping. Space assets are current limit on DOD launch capacity.1 In- typically designed to conform to a particu- creasing this capacity necessitates either lar launch vehicle’s limitations (e.g., engi- larger launch vehicles or higher perfor- neers might design a satellite to be carried mance from existing ones. Larger vehicles by a Delta IV-2 medium launch vehicle). drive a series of additional expenses, includ- Essentially, this choice of vehicle fixes the ing more propellant, expanded launch fa- maximum mass of the satellite and, thus, cilities, and bigger processing facilities. Al- its capabilities. If a launcher capable of though improved vehicles entail new placing more mass in the desired orbit were development costs, they may be compatible available at similar cost, the satellite’s de- with existing facilities. sign could allow for additional capability. Launching any medium or heavy ve- Furthermore, some payloads are too heavy hicle costs hundreds of millions of dollars. for present-day launch vehicles to place One estimate puts total launch costs of a into a particular orbit. A better-performing Delta IV Heavy launcher at $350 million; launcher would enable us to put those pay- other estimates are somewhat lower.2 A

*Lieutenant Colonel Hartsfield is an assistant professor of aerospace engineering at the Air Force Institute of Technology (AFIT), Wright-Patterson AFB, Ohio. Lieutenant Colonel Branam taught at AFIT for five years in the areas of combustion and jet and rocket propulsion. Currently, he is assigned as an instructor at the Air War College, Maxwell AFB, Alabama. Captain Hall is a master of science degree student at AFIT. Mr. Simmons is a PhD student at AFIT, studying space systems engineering.

Summer 2011 | 65 study by the RAND Corporation in 2006 Current Research: places launch costs for DOD payloads at $100–$200 million.3 The true expenditure of Improved Upper-Stage each launch is probably closer to the higher Current research at AFIT involves de- values at our current launch rates; however, signing and optimizing a cryogenic liquid more launches would push the cost per / upper-stage en- launch towards the lower values. Regard- gine. This new engine design, known as the less, launch expenses are immense. Using dual-expander aerospike nozzle (DEAN), the capacities and costs above, we can de- will serve as an orbit-transfer engine to termine that the price of lifting payload to propel a payload from LEO to GEO. The GEO amounts to $7,000–$25,000 per , DEAN differs from other cryogenic upper- and to LEO $2,000–$7,000 per pound. A stage in two ways. First, it utilizes Delta IV Heavy weighs about 1.6 million separate expander cycles for the oxidizer pounds at liftoff. Approximately 85 percent and fuel. Second, unlike bell-nozzle engines, (1.3 million pounds) is propellant (fuel and it employs an aerospike (radial inflow plug) oxidizer). If we assume an expenditure of nozzle (fig. 1). approximately $5 per pound for both hydro- In a typical engine-, the gen and oxygen (averaged among hydrogen fuel alone regeneratively cools the combus- sources), then we spend about $6.5 million tion chamber and nozzle.5 Regardless of en- for propellant.4 Because the price of fuel gine design, the chamber walls require some depends upon the cost of natural gas (the form of cooling since combustion tempera- most convenient source of hydrogen), any tures typically reach about 5,000° F (stain- 6 estimates are quite volatile. However, even less steel melts at about 2,550° F). Energy substantial changes in the cost of hydrogen transferred to the fuel during regenerative will not have a great effect on overall ex- cooling acts as the sole driver for the turbo penses since the current propellant makes pumps that inject the fuel into the combus- up less than 5 percent of the overall launch tion chamber. Since the energy available to outlay; this simple analysis also applies to drive the pumps is limited to whatever heat transfer occurred during cooling, expander- the cost of oxidizer. Thus, two large catego- cycle engines typically have relatively low ries comprise about 95 percent of expendi- chamber pressures. Higher combustion- tures: launch base operations and launch chamber pressures would improve engine vehicle materials and production. Clearly, performance in three basic ways: First, reducing launch expenses entails (1) bring- greater pressures lead to more efficient ing down labor costs associated with the combustion and enhanced energy release launch base by using simpler processes and from the fuel. Second, higher pressures im- designing for maintainability and higher prove the potential specific pro- reliability, and (2) lessening material and duced by the engine—improving and labor expenditures associated with the ve- performance.7 Finally, elevated chamber hicle by making components reusable pressures lead to smaller chamber volumes where possible, simplifying assembly of the and potentially less engine , although launch vehicle, avoiding exotic materials, this advantage is partly offset by the in- simplifying the geometry of component creased material thickness necessary to parts to reduce difficult machining steps, withstand the greater pressure. and so forth. AFIT’s research in aerospike The RL-10, the standard evolved expend- rocket engines, sponsored by the Air Force launch vehicle’s upper-stage engine, Research Laboratory Propulsion Director- utilizes the expander cycle. This cycle has ate, seeks to increase vehicle performance the advantage of simplicity. Specifically, it and decrease launch costs. does not require the preburners or gas gen-

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Propellant In ow Flow Direction

Combustion Chamber

Aerospike Nozzle

Propellant In ow

Flow Direction Bell Nozzle Rocket Engine Combustion Chamber

Figure 1. Geometry of aerospike and bell-nozzle rocket engines erators needed by some other liquid-fuel pumps on the single shaft—one for fuel and cycles; it permits the use of lightweight one for oxidizer. Although seals separate fuel turbo pumps because the working fluids in in the turbine, fuel in the pump, and oxi- the turbines remain relatively cool (approxi- dizer in the pump, they have a potentially mately 80–440° F rather than 2,200–3,100° F disastrous failure mode. If a seal between seen in other designs), allowing designers the high-pressure fuel and high-pressure to choose lighter materials. Moreover, the oxidizer fails, the mixture of fuel and oxi- cycle facilitates smooth ignition and start- dizer can produce an explosion that would up because it reaches full thrust with a destroy the engine, launch vehicle, and pay- much more gradual ramp-up, whereas load. Separate fuel and oxidizer cycles have staged combustion and gas-generator cycles the advantage of physically separating the tend to yield full thrust very rapidly.8 oxidizer and fuel until injection into the Although the DEAN uses the expander combustion chamber, thus eliminating the cycle, it is unique in that the oxidizer and risk of explosions caused by failure of the fuel pass through separate expander cycles. interpropellant seals. Since the latter sce- The oxidizer cycle drives the oxidizer turbo nario represents one of the more cata- pumps, and the fuel cycle drives the fuel strophic failure modes in traditional turbo pumps. Since the pump and turbine ­expander-cycle engines, the DEAN’s dual- sides of turbo pumps must share a common expander design can improve operational shaft, seals separate the high-pressure safety and mission assurance.9 (pump) side and the low-pressure (turbine) The DEAN also uses a radial inflow plug side. A conventional expander-cycle engine nozzle primarily to enable the dual-expander has one turbine, driven by the fuel and two cycle but also to allow a shorter, lighter en-

Summer 2011 | 67 gine. The direct performance advantages of the spike is truncated or chopped short of the aerospike nozzle are not exploited in reaching a fine point, leaving a planar, the upper-stage application for which the blunt end (fig. 2). Research shows only neg- DEAN is designed. In low ambient pressure, ligible performance losses for the aerospike which applies to upper-stage engines oper- nozzle due to moderate spike truncation.11 ating at high altitudes, aerospike nozzles behave like conventional bell nozzles. For these missions, the rocket engine requires a DEAN Advantages and Design high expansion ratio for the nozzle, which Considerations increases the length and weight of the engine. For example, the Delta IV’s second-stage The DEAN design offers many benefits RL-10B2 engine has a deployable nozzle ex- over the currently operational orbit-transfer tension to attain the required expansion ra- RL-10B2 engine, all of which would save the tio; the extendable portion of the nozzle, Air Force money, improve mission assur- almost 6.5 feet long, weighs a little more than ance, and help assure access to space for 203 pounds (an additional 86 pounds of years to come. The DEAN engine, designed equipment supports deployment).10 In low for high performance, saves engine weight ambient pressure, the aerospike offers sav- and fuel, lends itself to manufacturing that ings in weight and size compared to an equiva- uses today’s technology, features robustness lent expansion-ratio bell nozzle, especially if and tolerance of extensive ground testing,

Propellant In ow Flow Direction

Combustion Chamber

Aerospike Nozzle Rocket Engine

Propellant In ow Flow Direction

Combustion Chamber

Aerospike Nozzle Rocket Engine (Truncated)

Figure 2. Geometry of truncated and nontruncated aerospike engines

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and incorporates features that eliminate amount of thrust, especially when the aero- some catastrophic failure modes for upper- spike is truncated, the interstage structure stage engines. can be made smaller and lighter compared Any design strives to improve upon pre- to the interstage for the RL-10B2. Doing so vious designs. Delta IV’s RL-10B2 represents equates to indirect benefits to the booster the current state of the art in upper-stage stage in weight, size, and performance. rocket engines, but the DEAN is designed to The considerations discussed above in- outperform that technology. When com- fluence the DEAN’s design. Its combustion pleted, AFIT’s current models indicate that chamber and nozzle will use standard metals the DEAN will provide just over twice the and ceramics compatible with the propel- thrust and weigh approximately 20 percent lants. Furthermore, the engine will use cur- less than the RL-10B2.12 Using a higher rent off-the-shelf turbo pumps and plumbing. ­propellant-mixture ratio (i.e., less fuel and Combined, these two features will improve more oxidizer), the DEAN will operate the design’s near-term manufacturability. leaner, demand less fuel, and thus decrease The DEAN’s designers wish to make the the money spent on fuel slightly since liq- engine reusable and robust enough to with- uid oxygen is somewhat cheaper than liquid stand extended ground testing prior to hydrogen. Furthermore, AFIT performance launch. Taking a conservative approach, calculations indicate that matching or im- AFIT engineers determined a maximum proving the of the RL-10B2 wall temperature for both the combustion results in a minimum stage-weight savings chamber and aerospike that would prevent of 105 pounds due to the reduced estimated degradation of material strength. Our mod- weight of the DEAN.13 Any improvements eling rejected designs unable to maintain in specific impulse would enable additional combustion chamber and aerospike tem- weight savings for the launch vehicle as a peratures below the limits established for whole. The higher the specific impulse, the the materials simulated. less propellant needed to realize the desired thrust. This weight savings permits an in- crease in payload weight, which may include Future Work: the addition of new capabilities to the satel- High-Performance Booster Engine lite being launched. Because of the costliness of launches, a savings in weight equates di- The step in aerospike rocket re- rectly to one in expenditures; therefore, a search at AFIT calls for applying the aero- 105-pound weight savings can save the gov- spike nozzle to first-stage (booster) engines. ernment on the order of $1 million (at The nozzle offers the significant perfor- about $10,000 per pound, based on mean mance advantage of operating nearly opti- values of the costs discussed earlier).14 mally at all altitudes below its design alti- Utilizing an aerospike upper stage also tude, thanks to a capability known as brings indirect benefits to the first-stage altitude compensation. Conversely, a con- booster. The interstage (part of the first ventional bell-nozzle engine, such as the stage) encapsulates the upper stage to pro- ’s main engine, is designed for tect its components during atmospheric optimal operation at a single design alti- travel. This component is dead weight in tude, suffering performance losses at all the sense that, though necessary for the other altitudes. The aerospike design has mission, its weight decreases the amount of significant performance advantages during payload, engine, and propellant the vehicle operation through the atmosphere. In can carry, so engineers seek to make the rocket engines, the nozzle expansion ratio interstage as small and light as possible. Be- is a key to maximizing engine performance. cause the aerospike design is shorter than a A high expansion ratio leads to low exhaust bell nozzle and can produce the same pressure, increasing the conversion of po-

Summer 2011 | 69 tential output (represented by the chamber from the fact that the outside diameter of temperature and pressure) to thrust output the engine effectively sets that ratio; thus, (exhaust momentum and pressure). Ex- an extremely large expansion ratio requires haust pressures in excess of the ambient a very large-diameter engine, adding con- atmospheric pressure for the flight altitude siderable weight. The challenge lies in bal- generate some thrust, but a larger expan- ancing the increased performance with the sion ratio could convert that extra pressure increased weight to find an optimal point into increased momentum and more thrust for the launch vehicle. than the pressure alone can provide. There- This near-ideal performance becomes fore, for all rockets, the largest expansion especially important during the low-altitude ratio nozzle possible represents a perfor- boost phase of the rocket flight. With no mance advantage. However, for conven- other performance changes to the launch tional bell-nozzle rocket engines, the nozzle’s vehicle, AFIT’s initial modeling studies indi- size has limitations. If the exhaust pressure cate that changing the first-stage engine to is less than about 25–40 percent of the am- aerospike nozzle engines could produce an bient pressure, the exhaust flow will sepa- approximately 6 percent increase in the rate within the nozzle, forming shock waves mass that the vehicle can lift to GEO. The and causing large thrust losses. To avoid difference in performance, calculated for this condition, engineers generally design identical chamber pressures and mixture rocket engines to operate with exit pres- ratios, could see improvement with changes sures no lower than about 60 percent of the to these and other parameters. AFIT’s re- ambient pressure, providing some margin search aims at identifying an optimal en- of safety.15 This sets a practical limit for gine design (or a set of optimal designs) bell-nozzle expansion ratio, based on the that may not share operating conditions lowest altitude at which the rocket is ex- with current lift engines such as the RS-68 pected to operate. Normally, the engine de- used in the Delta IV launcher. Performance signer sets the design altitude to about alterations such as increasing the combustion- 12,000 feet, where the atmospheric pressure chamber pressure can significantly enhance is about 62 percent of sea-level pressure.16 specific impulse and payload capacity. If Setting the design altitude any higher cre- the aerospike operates at double the RS-68’s ates the potential for separated flow within chamber pressure, the improvement in the nozzle and greatly reduced thrust. mean specific impulse also doubles, as does Therefore, at all altitudes above that, the the increase in payload capacity to GEO. rocket produces substantially less thrust We have modeled the performance of a than it could ideally (see fig. 3). conventional bell-nozzle rocket, an aerospike- The aerospike nozzle does not suffer nozzle rocket, and an ideal rocket with an from this disadvantage. Increased ambient infinitely adjustable area-ratio nozzle and pressure effectively reduces the expansion no thrust losses due to friction or other fac- ratio to a point where the exhaust pressure tors (fig. 3). The conventional rocket, built matches the ambient pressure. In this way, around a 12,000-foot design altitude to allow the aerospike nozzle compensates for alti- separation-free operation at sea level for tude up to its design altitude, represented launch, assumes a 95-percent-efficient noz- by its physical expansion ratio. Above this zle design to account for friction and other altitude, the aerospike nozzle acts much loss effects. Note that the specific impulse like a bell nozzle, with the excess exhaust remains below that of the aerospike at all pressure generating some extra thrust as altitudes except 12,000 feet. Furthermore, the rocket climbs above its design altitude. the shape of the curve for the conventional Since no fluid-dynamic reason exists for rocket does not track the ideal nozzle, indi- limiting the nozzle expansion ratio, the cating less-than-optimum performance at practical limit to the aerospike’s ratio comes all altitudes. The aerospike rocket features

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Speci c Impulse Variation with Altitude

490

470

450

430

410

390 Speci c Impulse (seconds)

370

350 0 20 40 60 80 100 120

Altitude (thousands of feet)

Aerospike Bell Nozzle

Aerospike Design Point Bell Nozzle Design Point Ideal Engine

Figure 3. Performance advantage of aerospike engines in the atmosphere chamber conditions identical to those of the expand from about six feet in diameter at conventional rocket but has a design alti- sea level to almost 52 feet in diameter at tude of 43,000 feet since that setting pro- 118,000 feet. To continue this performance duced an engine slightly smaller than the until the rocket reaches near vacuum at diameter of a Delta IV first stage. The figure 262,000 feet, the nozzle would have to ex- shows that the aerospike’s specific-impulse pand to 672 feet in diameter—clearly im- curve runs parallel to the ideal curve, up to practical. Long before this point, the engine 43,000 feet. The aerospike curve assumes a would become too heavy to lift itself, much 95-percent-efficient nozzle to account for less any fuel or payload. losses, thus falling below the ideal. Notably, Through a boost of slightly more than 3 although the aerospike nozzle has a diameter percent in mean specific impulse on the of nearly 13 feet to reach exhaust-gas ex- first stage with an aerospike, without ac- pansion appropriate for pressure conditions counting for any weight savings by using at 43,000 feet, the adjustable nozzle must the DEAN engine on the upper stage(s),

Summer 2011 | 71 current AFIT modeling indicates the possi- Conclusion bility of realizing a 6 percent gain in maxi- mum payload to GEO. Improving from a As an Air Force, we find ourselves at a Delta IV payload limit of 14,491 pounds to decision point for space operations. Most of GEO to 15,355 pounds would enable a sig- our rocket engines reflect decades-old tech- nificant increase in spacecraft capability as nology in all aspects of their construction. well as a decrease in the payload’s launch Costs are high, and the vehicles are gener- cost per pound. Doubling the chamber pres- ally not reusable, even if we recover them sure produces a 6 percent rise in specific after launch. At AFIT, our rocket team impulse and a 13 percent increase in GEO thinks that the Air Force can do better. The payload—to 16,437 pounds. Similar perfor- reduced weight of the DEAN would result mance improvements would also result from utilizing the first-stage aerospike en- in incremental improvements to launch ca- gine to attain LEO orbits. pacity without extensive reworking of the As with the DEAN’s upper-stage engine, lower stages. The increased specific im- the aerospike-nozzle booster engines would pulse available from the aerospike first- be more compact than conventional bell- stage engine could produce a significant im- nozzle engines. Replacing the bell nozzle provement in the satellite weight we can with the radial-inflow plug nozzle can ex- place in orbit. Currently, the overall weight pand the maximum diameter of the engine, of the launch vehicle limits the capabilities but using a truncated aerospike allows a of our space platforms. In many cases, we much shorter engine. Doing so can trans- must omit adjunct payloads that could offer late into weight savings and might make the new or enhanced capabilities because we aerospike engines more adaptable to multi- simply cannot launch the extra weight or engine operations for larger lift capabilities. provide electrical power (more power im- AFIT set a goal of improving perfor- plies more weight in solar panels) to sup- mance and producing a more compact en- port the additional equipment. Enhancing gine while maintaining operability with key subsystems such as propellant pumps our launch capability helps solve this prob- and materials. By ensuring that the perfor- lem. Moreover, designing engines for reli- mance required of the turbo pumps lies ability, maintainability, and operability within that demonstrated in testing for re- from the start will improve launch costs alistic launch conditions (the National and launch rates. At AFIT we believe that Aeronautics and Space Administration re- the Air Force needs a push in the direction fers to this as technology readiness level of building an updated launcher since we six, a system adopted by the DOD acquisi- know that developing the technology will tion community), AFIT can reduce the take many years, and building a new risks associated with depending on outside launcher many more years. As an air and 17 developments. By restricting material space force, we cannot wait for obsoles- choices to conventional metals and ceram- cence of current platforms to start develop- ics, the AFIT design team can avoid need- ment of a follow-on space launch platform. ing any breakthroughs in materials. How- We must start now, and AFIT research is ever, the team will take advantage of any pointing the way.  such advancements in scientific material to further improve the aerospike engine’s Wright-Patterson AFB, Ohio performance in the future. Maxwell AFB, Alabama

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Notes

1. United Launch Alliance, Delta IV Payload 9. 2nd Lt David F. Martin II, “Computational De- Planners Guide (Littleton, CO: United Launch Alli- sign of Upperstage Chamber, Aerospike, and Cooling ance, 2007), 2-10. Jacket for Dual-Expander Rocket Engine” (master’s 2. Todd Halvorson, “Rocket Analysis: A Look at thesis, Air Force Institute of Technology, March What Could Sway Obama’s Decision on a New 2008), 19, https://www.afresearch.org/skins/rims Launch Vehicle for NASA,” Florida Today, 8 February /q_mod_be0e99f3-fc56-4ccb-8dfe-670c0822a153/q_act 2009, http://www.floridatoday.com/assets/pdf _downloadpaper/q_obj_8ace2da5-a079-4a4b-91d3 /A912809629.PDF. -7296764e676c/display.aspx?rs=enginespage. 3. RAND Corporation, National Security Space 10. George P. Sutton, History of Liquid Propellant Launch Report (Santa Monica, CA: RAND Corporation, Rocket Engines (Reston, VA: American Institute of 2006), 32, http://www.rand.org/pubs/monographs Aeronautics and Astronautics, 2006), 496–97; and M. /2006/RAND_MG503.pdf. Lacoste et al., “Carbon/Carbon Extendible Nozzles,” 4. Dale R. Simbeck and Elaine Chang, Hydrogen Supply: Cost Estimate for Hydrogen Pathways—Scop- Acta Astronautica 50, no. 6 (March 2002): 357–67. ing Analysis, January 22, 2002–July 22, 2002 (Moun- 11. Takashi Ito, Kozo Fujii, and A. Koich Hayashi, tain View, CA: SFA Pacific, November 2002), 12, “Computations of Axisymmetric Plug-Nozzle Flow- http://www.nrel.gov/docs/fy03osti/32525.pdf. fields: Flow Structures and Thrust Performance,” 5. Regenerative cooling uses propellant flow AIAA Paper 99-3211 (Reston, VA: American Institute through passages surrounding the combustion of Aeronautics and Astronautics, 1999). chamber and nozzle to reduce temperatures in the 12. J. Simmons and Richard Branam, “Paramet- chamber and nozzle walls. After flowing through ric Study of Dual-Expander Aerospike Nozzle Upper these passages, the propellant is injected into the Stage Rocket Engine,” Journal of Spacecraft and Rock- combustion chamber. Heat absorbed from the cham- ets (forthcoming). ber is returned to the chamber and produces higher 13. Ibid. temperature combustion. Dieter K. Huzel and David 14. Halvorson, “Rocket Analysis”; and RAND Cor- H. Huang, Modern Engineering for Design of Liquid- poration, Space Launch Report, 32. Propellant Rocket Engines (Washington, DC: American 15. Sutton and Biblarz, Rocket Propulsion Ele- Institute of Aeronautics and Astronautics, 1992), 84. ments, 70. 6. George P. Sutton and Oscar Biblarz, Rocket 16. John D. Anderson Jr., Introduction to Flight, Propulsion Elements, 7th ed. (New York: John Wiley 6th ed. (Boston: McGraw-Hill, 2008), 847. & Sons, 2001), 188; and Frank P. Incropera et al., 17. Technology readiness levels (TRL) run from Fundamentals of Heat and Mass Transfer, 6th ed. one to nine, with one essentially an idea and nine (New York: John Wiley & Sons, 2006), 931. 7. Specific impulse is the ratio of thrust to weight designating technology in operational use. TRL six flow of propellant through the engine. Large num- represents technology considered ready for opera- bers are better and imply that less propellant is re- tional demonstration, having been proven in repre- quired to carry out any specific mission. Sutton and sentative testing. Defense Acquisition University, Biblarz, Rocket Propulsion Elements, 28. Defense Acquisition Guidebook (Ft. Belvoir, VA: De- 8. Detlef Manski et al., “Cycles for Earth-to-Orbit fense Acquisition University, 16 February 2011), Propulsion,” Journal of Propulsion and Power 14, no. chap. 10, sec. 5.2.2, https://acc.dau.mil/adl/en-US 5 (September–October 1998): 588–90, 594. /350719/file/49150/DAG_02-16-11.pdf.

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