High Pressure Bipropellant Engine System Study

Ron Portz*, David Krismer†, Frank Lu‡ and Scott Miller§

Aerojet General Corporation, Redmond, Washington, 98073, Log Number 2006-009

A trade study has been completed to evaluate the spacecraft-level performance increase attainable by using high pressure bipropellant engine technology. Figures of merit at the spacecraft level include propulsion subsystem wet mass and volume, power required, pressurization system reliability/complexity, and overall cost. Missions studied included geostationary satellite apogee maneuvers and demanding planetary spacecraft delta-V maneuvers such as Mercury Messenger, Europa Orbit Insertion and Titan Orbit Insertion. An engine performance model was created by using TDK software anchored to Aerojet’s flight heritage High Performance Apogee Thruster (HiPAT™) data. TDK was run over a wide range of parameters to obtain predictions for a regression-based performance analysis. The parameters included:

• Chamber pressure (150-700 psia) • Thrust (100-lbf to 500-lbf) • Mixture ratio (0.8 to 1.5) • Area ratio (200:1 to 500:1) • Propellants (N2H4 and MON-3, MON-10 or MON-25) • Radiation or regenerative cooling

Performance regression equations were then inserted into a spacecraft performance model to determine the net wet mass savings over a range of operational points. The impact of higher chamber pressure on engine size, mass and cooling is addressed. Based on the results of the study, recommendations are made for near term technology development and risk reduction efforts in order to implement this system approach.

Nomenclature AR = Area Ratio FFC = Fuel Film Cooling GEO = Geosynchronous Earth Orbit hg = Convection coefficient Isp = Specific impulse LAE = Liquid apogee Engine MMH = Monomethylhydrazine, N2H3CH3MON-X = Mixed oxides of nitrogen, nitrogen tetroxide and X% NO by mass in solution N2H4 = Hydrazine, N2H4NTO = Nitrogen tetroxide, N2O4OF = Oxidizer to Fuel ratio Pc = Chamber pressure psia = Pounds per square inch absolute SG = Specific Gravity SOA = State of the Art TDK = Two Dimensional Kinetic

* Sr. Project Engineer, Systems and Technology Development † Engineering Specialist, Systems and Technology Development, AIAA Senior Member ‡ Engineering Specialist, Systems and Technology Development, AIAA Member § Manager of Programs, Systems and Technology Development, AIAA Senior Member

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42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit9 - 12 July 2006, Sacramento, California

AIAA 2006-5219

Copyright © 2006 by Aerojet-General Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

I. Introduction ERap

storage

OJET has an interest in improving the performance of liquid in-space propulsion, both for Earth orbit plications and exploration missions. Despite known safety issues, existing infrastructure and long-term strongly favor conventional Earth-storable propellants, such as nitrogen tetroxide (NTO, N2O4) and either

monomethylhydrazine (MMH, N2H3CH3) or the higher-performing hydrazine (N2H4). Fluorinated oxidizers and metallized fuels offer performance improvements1, but at unacceptable costs for both environmental compliance and new facilities. Liquid oxygen as an oxidizer presents several advantages in terms of reduced toxicity, increased performance and existing infrastructure, but the required long-term, subcritical storage capability has yet to be demonstrated.

A

With few exceptions, in-space propulsion, as distinct from Earth-to-orbit propulsion, employs pressure-fed engines with Earth-storable propellants. This propulsion technology is mature, extremely reliable and well-studied, however the pressure-fed condition imposes some performance limits on an optimized system. High thrust in a compact engine and to a degree, high specific impulse, are typically associated with high chamber pressure, Pc, but simply increasing tank pressure and adding pressurizing gas result in high weight that may result in a net decrease, rather than increase in performance of the space vehicle or incompatibility with the desired launch vehicle.

Pressure-fed rocket engine nozzles are frequently radiation-cooled with very large area ratios, AR, approaching 400, to maximize performance. The dimensions of these engines often present a limit on vehicle envelope due both to the physical size of the long nozzle and the need for unimpeded field-of-view to space. A strategy to increase engine and system performance and manage engine size is to pump the propellants to high pressure prior to chamber injection. The subject of pumps for small storable propellants is addressed by a handful of researchers at most space and propulsion conferences e.g. 2, 3.

This paper will report on a trade study conducted by Aerojet to quantify the performance potential of pumped engines and translate the performance increase to system benefits for a number of real and hypothetical space missions. The impact of the pump on system reliability and engine cooling as design and cost drivers will be addressed and the state of the art for pump design and development for Earth-storable propellants will be reviewed.

II. TDK Performance Predictions Anchored to Test Data The TDK computer program4 was used to predict engine performance, with efficiency factors used to correlate

results with Aerojet test data. TDK is a commercially available computer code that implements the standard JANNAF methods for predicting rocket nozzle performance. The performance of hypothetical rocket engines was predicted over chamber pressures from 300 to 700 psia, thrust from 300 to 500 pounds, oxidizer to fuel ratios of 0.8 to 1.4 and AR from 200 to 500. The baseline engine used for comparison is Aerojet’s Dual-Mode High Performance Apogee Thruster, or HiPAT shown in Fig. 1, which nominally delivers 445 N (100 pounds) of thrust at an OF ratio of 0.85 and AR of 3755. The baseline propellants used in this analysis were hydrazine, N2H4, and the mixed oxides of nitrogen oxidizer MON-3. This combination represents the high-performance state-of-the-art in Earth-storable liquid propellants and makes available the dual-mode main-propulsion/ACS option exercised by many spacecraft designers to minimize system mass by allowing the hydrazine to be used either as a high performance monopropellant or bipropellant fuel. Some performance enhancement is predicted upon using MON-10 and MON-25 to oxidize the fuel and these alternatives were also evaluated, recognizing that there are availability and combustion stability issues to be dealt with separately.

It is essential to anchor the TDK predictions to real hardware test data. TDK allows the user to estimate nozzle efficiency, with losses coming from boundary layer formation and growth, but does not account for combustion efficiency. This is done by performing a TDK analysis at “real-life” HiPAT operating conditions and comparing the

Figure 1. Aerojet R-4D-15 HiPAT during hot-fire testing.

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analytical prediction with the recorded performance of the hardware. An estimate of obtainable efficiency is obtained and applied to the hypothetical operating conditions.

Comparing HiPAT test data with TDK data in Fig. 2 shows a significant difference from the OF ratio resulting in maximum HiPAT Isp. The main reason for this difference is the use of fuel film cooling (FFC). About 30% of the total fuel flow is injected near the inner chamber wall to chemically and thermally protect the wall from the core

flow, which consists of fuel and oxidizer mixed in near stoichiometric proportions. With a given injector design, the overall O/F ratio can be varied by altering feed pressures, but the FFC percentage will remain fixed. Another result of FFC is that the core flow is diluted with raw fuel near the wall, reducing performance. In Fig. 2, these effects are represented by dividing the OF ratio corresponding to a prediction point by 1.17. With 30% FFC and no mixing, this number should be 1.3, but enhanced mixing is a trait of the HiPAT, so variation is expected. Reduced efficiency compared to the TDK prediction is accounted for by a 97% efficiency factor resulting in the final, “Eff. Adjusted TDK Isp”

line in Fig. 2. This result is partially accounted for by the fact that during development, the dual mode HiPAT nozzle was not optimized for the higher performing propellant combination in the interest of using existing tooling, leaving ~0.5% performance increase available.

An additional reason for improved performance in the TDK data following compared to HiPAT data is increased engine thrust. At a given combination of propellants, chamber pressure and expansion ratio, higher thrust is achieved by physical scaling of the engine. Larger physical size of the chamber and nozzle means that the viscous boundary layer occupies a smaller volume of the thrust chamber cross-section. The nozzle exerts proportionally less drag on the flow and slightly higher Isp results. Increasing thrust from 100 to 500 pounds resulted in approximately 0.5% Isp increase attributable to the thrust increase. Between the known optimization of the HiPAT, and improved

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Figure 2. Comparison of TDK predicted HiPAT performance with test.

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Figure 3. Isp versus OF ratio at 500lb thrust for N2H4/MON-3 propellants in a radiation cooled engine

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combustion efficiency due to eliminating FFC an efficiency factor of 98% is applied to TDK data following to derive performance estimates.

The higher performance options presented following do not assume any FFC, due both to limitation of the TDK program and the desire to maximize performance. A final engine design would incorporate some degree of FFC, but at a reduced level. The reduction in Isp is accounted for in the charts and the effect on the OF relationship to Isp will be minor. System performance is enhanced both by the increased chamber temperature resulting from elimination of FFC, and by the increased density impulse resulting from burning a higher fraction of oxidizer (SG = 1.44) in the engine versus fuel (SG = 1.01). The additional materials engineering challenge presented by reducing or eliminating FFC must be addressed separately and weighed in the development cost decision process.

The calculated results shown in Fig. 2 used the assumption of radiation cooling for consistent comparison with HiPAT test data. This assumption requires energy to radiate from the chamber, as shown very graphically in Fig. 1, and is lost to the engine. Using the propellant to regeneratively cool the chamber and nozzle returns a significant fraction of this energy to the engine, resulting in decreased system entropy loss and improved Isp

6. Fig. 4 shows that TDK-predicted Isp increases 2 to 3 seconds when regeneration is employed. The increase predicted by applying the entire heat radiated to increasing propellant exit velocity is higher, but is not realized because most of the engine heat is lost through the nozzle extension, due to its large surface area even though the temperature is much lower than the chamber and throat. Practical considerations of cooling channel size and manufacturability do not lend to cooling the entire nozzle for small, high-expansion ratio engines. The simulation used regenerative cooling to an AR of 7 with a radiation cooled extension.

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Figure 4. Isp vs OF ratio at 500 lb thrust for N2H4/MON-3 propellants in a regeneratively cooled engine.

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Figure 5. Relative effect of NO Percentage in MON-X Oxidizer

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Use of an oxidizer other than MON-3 (3% NO in N2O4 solution) results in TDK predicting the variation in Isp shown in Fig. 5, in agreement with experimental data7. The Isp improvement is significant and the increasing optimal OF improves system performance despite the slightly lower densities of higher MON oxidizers. This improvement comes with a price. Oxidizer with NO percentage higher than 3 is not readily available. Test programs using this propellant begin with the tedious process of making a batch7. There are also believed to be engine stability issues that must be understood and mitigated, although none are noted in reference 7. On the other hand, the lower freezing temperature of higher NO percentage oxidizers is seen as a bonus for interplanetary missions, as it decreases the vehicle power demand for heating propellant tanks and lines to prevent freezing. In aggregate, the gains are not felt to justify the lack of availability so only MON-3 is considered in system comparisons.

III. Other Aspects of Engine Design

A. Engine Size and Weight Increasing Pc has a beneficial effect on engine weight.

Comparing two geometrically similar engines with the same thrust and propellants, but differing in Pc, the throat area required varies approximately inversely with the pressure so the total engine surface area also varies with the inverse of pressure. Wall thickness must increase to withstand the higher pressure, but by the square root of the pressure increase, due to the reduced engine diameter. The result is that the engine weight varies with the inverse square root of the pressure increase. For example, increasing Pc 9 times results in reducing engine weight by a factor of three. Examination of Fig. 6, at any line of constant thrust reveals this trend

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Using pumps to increase Pc will add mass in the form of pumps and power conditioning components. For system modeling purposes, a linear formula based on Aerojet empirical data has been used to estimate mass of these components.

Figure 6. Predicted variation of chamber and nozzle weight with chamber pressure and thrust. Note that practical considerations of wall thickness minimize this impact. Increasing thrust while holding Pc constant results in a

predicted weight increase. Surface area increases with thrust, and the wall thickness must increase with the square root of thrust to maintain constant stress, so engine material volume, hence mass, increases with thrust raised to the 3/2 power. This theoretical trend is also visible in Fig. 6, for any line of constant Pc.

In practice this weight increase is not realized, since low pressure and thrust engines are not optimized for wall thickness. They wind up with wall thicknesses much greater than pressure forces demand due to the need for manufacturability, to resist handling, and to survive launch dynamic environments. For predicting engine weight in a system model, the increase in wall thickness compared to the baseline is ignored, assuming this thickness has been fixed for manufacturing, handling, and structural vibration reasons. Engine weight then varies directly with thrust and inversely with the square root of Pc.

B. Pump Power Pumps require a power source. Pump driven engines

usually derive their power from coolant expansion, a gas generator or a preburner, which emit gas to spin a turbine powering centrifugal compressors8, 9. Small, in-space engines may take advantage of the underutilized electrical

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Figure 7. Size constrained operating space for pumped in-space liquid engine.

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power capacity of the overall vehicle during periods of otherwise low power consumption10. A great advantage of this arrangement is that no propellant is consumed which would decrease the delivered Isp. However, the power limit constrains engine chamber pressure and thrust.

Engine physical size and radiation cooling are design constraints for many Aerojet customers. Figure 7 shows that this constraint limits the thrust and expansion ratio that can be employed at a given value of Pc. If the additional constraint of limited power to drive the pumps is imposed, then the operating space decreases in size further, as shown in Fig. 8 for a hypothetical 5 kW power limit. Since Isp is strongly dependent on Pc and expansion ratio, these constraints can limit engine and system performance.

The extent of the power-limited operating space is affected strongly by the efficiency of the pumps and power conditioning equipment. 50% pump efficiency, and 92% electric-to-mechanical power conversion efficiency were used to generate Fig. 8 and in system modeling, and while these values are based on considerable experience of specialists at Aerojet, they would benefit from validation with test of flight configuration hardware.

Increased pump efficiency would permit higher Pc with increased Isp and reduced engine size for a given power availability. For system modeling, it is assumed that electrical power is proportional to vehicle launch mass.

C. Cooling As Pc and thrust both increase, the viability of building a regeneratively cooled chamber increases. This is

important because the rhenium and iridium used in the chamber of the SOA, radiation-cooled LAE constitute approximately 64% of the total materials cost and 30% of the total selling price of a single engine. If common, easily formed materials can be used in a regeneratively cooled chamber, the cost reduction may offset the cost of pumps and the power control hardware necessary to operate them. Regenerative cooling may aid packaging and the results presented in Fig. 4 also show an Isp increase of approximately 0.8% for regenerative cooling of the chamber and a portion of the nozzle.

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Figure 8. Size and 5 kW power constrained operating space for pumped in-space liquid engine.

The rate of heat transfer per unit area from the combustion chamber gas into the chamber wall is expressed by9: ( )whwag TThq −= *" (1) Where Twh is the temperature of the gas-side, or hot, wall. Assume that Twa, the adiabatic wall temperature, is equal to the fluid stagnation temperature, T0. This is nearly true at the nozzle throat, but in other parts of the nozzle, T0 is higher than Twa, so this assumption yields a conservative heat transfer calculation. Further assume that Twh is equal to the “cold-side” wall temperature (coolant temperature for a regeneratively cooled engine). For a thin, metallic wall, this assumption is usually good because of high thermal conductivity. If propellant is used to cool the engine, Twh must remain in a relatively narrow range for coolant to remain liquid in the nucleate boiling regime (about Tsat + 50)11, so Twh is constant relative to the hot gas, for practical purposes.

An empirical relation for hg is given by9

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b

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cGDGc

h μμ , (2)

where kb is thermal conductivity, μb is viscosity, specific heat is cp, and average mass flow per unit area is G. The subscript b indicates bulk properties. This can be re-written as . (3) 67.02.047.033.08.0 *****023.0 bbpg kDcGh −−= μ

Comparing the effect of Pc with constant chamber size, and recognizing that Twa and the dependent properties cp, μb, and k b do not change much with Pc, the convection coefficient can be compared between two alternatives as . (4) 8.0Ghg ∝

G is the mass flow per unit area, or ρu,

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( ) 8.0uhg ρ∝ . (5) For geometrically similar nozzles, the velocity at

corresponding points is equal, so the convection coefficient varies as density, or for perfect gases, the familiar result that heat transfer varies as Pc to the 0.8 power9. Increasing pressure results in higher heat flux, but the propellant flow increases approximately linearly to maintain the higher pressure. As a result, the FFC percentage or heat capacity required to regeneratively cool the chamber decreases with higher Pc. 1

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Figure 9. Normalized variation of convection coefficient with chamber pressure and engine diameter.

For two engines with equal propellants, OF and Pc, but different size, the fluid stagnation temperature, hence Twa, will be equal. Twh will also be equal for the two engines, because the coolant temperature will be the same. So for two engines, differing only in size (thrust), the change in heat transfer per unit area is primarily a function of hg. Referring to Eq.(3)

67.02.047.033.08.0 *****023.0 bbpg kDcGh −−= μ For the two engines, kb, μb, cp, and G, are approximately equal. Therefore, . (6) 2.0−∝ Dhg

Physically large engines exhibit lower heat flux, hence are easier to cool than small engines, given the same propellants, Pc, etc. Considering both Pc and size shows that for constant thrust, where Pc increases lead to size decrease, the convection coefficient increases as, . (7) 9.0

cg Ph ∝which is not much help at all. Engine thrust (size) must increase for cooling to improve significantly. Comparing the total heat rejection of a hypothetical engine to a reference engine of similar geometry,

propellants etc., but different size, the total rate of heat transfer is approximately

refrefc

c

refQP

PA

AQ **8.0

,⎟⎠⎞

⎜⎝⎛= (8)

Where A is the surface area of concern, and Q is the total heat flow. The subscript ref signifies a reference engine, for which the total, radiation heat flow has been determined, based on surface temperature and emissivity. The surface area of a hypothetical engine is proportional to the throat area, which is roughly proportional to the thrust, F and inversely proportional to the chamber pressure, so

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All of this heat must be absorbed by the coolant, so assuming an inlet temperature, constant specific heat, and scaling the propellant mass flow rate with thrust and pressure, the coolant outlet temperature can be estimated. The heat rejected by a regeneratively cooled engine will be higher than a radiation cooled engine because of the greater temperature gradient both within the chamber walls and across the thermal boundary layer. This has been accounted for in the input for Fig. 10, which shows approximate coolant outlet temperature. As engine pressure and thrust rise, the average coolant outlet temperature becomes more manageable. The rate of heating must be given special care to avoid the film boiling regime both to maximize heat transfer and prevent thermal decomposition11. This is critical to ensure propellant stability. Reactive propellants like hydrazine in a regenerative cooling application demand special

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care during start-up and especially during shut-down transients to prevent explosive decomposition during thermal soakback. This is expected to require inert gas purge of the coolant lines at startup and shutdown, although tests have not exhibited problems12.

D. Reliability One of Aerojet’s founders, Dr. Theodore von Kármán

observed, It is my personal belief that the length of the period of attaining reasonable reliability in the development process could be essentially reduced if simple design were emphasized as a leading principle, even if we had to make some sacrifice in the quantitative measure of “efficiency.” Essential elements have to be designed as simply as possible, even if this means a reduction in quantitative efficiency and a certain increase of bulkiness and/or weight.8

Modern hypergolic, earth-storable rocket engines are extremely simple and reliable in operation, with remarkably few moving parts and exceptional service records. System failures are

probably more likely to be attributed to the storage and pressurization system than the rocket engine. While adding complexity in the form of pumps is counter to Dr. von Kármán’s advice, the impact on system reliability is minor. By adding pumps, the demand on propellant storage and pressurization systems is lessened. Quantization of the change in reliability is not seen as a credible exercise at this time given the paucity of data on small pump reliability.

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Figure 10. Coolant outlet temperature estimated based on engine thrust and chamber pressure.

IV. Reference Missions Engine performance has little meaning separate from system performance. To estimate the real effect on the

spacecraft, a math model of the propulsion system was created, wherein the masses of engines, propellants, pressurant tanks and ancillary components are calculated based on pressures and propellant needs. This exercise has the potential to devolve into fiction, so Aerojet has used real hardware masses to ground the exercise. When estimates are required, erring on the side of minimum mass was preferred since performance improvements had less of an effect on vehicle mass, so were conservative.

Four reference missions were identified for comparison of the impact of the improvement on system performance. They include a generic geosynchronous Earth satellite, the Messenger spacecraft currently on its way to Mercury and two hypothetical spacecraft: an orbiter for Jupiter’s moon Europa and an orbiter for Saturn’s moon Titan. These reference missions were chosen for the economic importance of GEO and the high demand placed on the propulsion system of the exploration missions. Aerojet is fortunate to have constructed the propulsion systems for these or similar vehicles, and so has reasonable data from which to estimate system mass and state-of-the-art performance. For the data presented in Table 1, baseline engine performance is anchored either to supplier data for Messenger’s Leros 1b engine or Aerojet data for the dual-mode HiPAT thruster with expansion ratio of 375 for all others. For the pumped system performance improvement, regeneratively cooled chambers with expansion ratios of 400 are assumed burning MON-3 and N2H4.

Practical considerations complicate quantifying the advantage to the spacecraft. As the required propellant mass and storage pressure decreases, the propellant and pressurant tank sizes and masses can potentially go down, however the availability of commercial, qualified tanks usually means that oversized, rather than optimized units are selected for use. The effect is that while mathematical scaling is useful to bound the eventual component mass, system cost will usually drive design to the next largest qualified unit that meets minimum requirements, which may well be the reference design.

The worth of the mass delta is subjective. For a revenue driven enterprise like a commercial GEO satellite, this cost is very relevant to increasing on-orbit lifespan or adding revenue-generating capacity, however for the science missions the value of increased performance to either enabling the mission or permitting inclusion of a desired experiment may be beyond easy quantification. For example, the Mercury Messenger spacecraft is carrying slightly more than 42 kg of instruments13. Use of a pumped propulsion system has the potential to nearly double the mass that could have been devoted to instruments by reducing the propellant need and decreasing propulsion system dry mass. For the outer planet moon missions the probable increase in vehicle capability will be more dramatic.

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Table 1. Reference missions and performance improvement

GEO Messenger Europa Orbiter

Titan Orbiter

Launch Vehicle Deployed Mass (kg) 4800 1093 639 1172 Spacecraft Propulsion Delta V (m/s) 1900 2081 5500 3455 Baseline Performance

Steady-State Isp (s) 327 318 327 327 Steady-State Thrust (N) 445 667 445 445 Delta-V Propellant Required (kg) 2146 532 524 773 ACS & Reserve Propellant (kg) 380 68 34 70 Total Propellant Load (kg) 2526 600 558 843 Pressurant Load (kg) 5.6 1.32 1.22 1.85 Propulsion System Mass (kg) 229 81.5 77.8 102 Delivered Mass (kg) 2654 561 115 399

Pumped System Performance Steady-State Isp (s) 344.3 341.7 341.7 341.7 Steady-State Thrust (N) 1423 890 890 890 Chamber Pressure (bar) 41.4 27.6 27.6 27.6 Electrical power available (kW) 5.0 2.0 2.0 2.0 Delta-V Propellant Required (kg) 2066 506 515 754 Propellant Load (kg) 2446 572 549 824 Pressurant Load (kg) 2.51 0.56 0.53 0.80 Propulsion System Mass (kg) 182.5 68.3 66.7 84.8 Delivered Mass (kg) 2783 601 135 435 System Mass Delta (kg) 129 39.8 20.1 36.3 Percentage Increase in delivered mass 4.9% 7.1% 17.5% 9.1%

Worth of Mass Delta @ $10,000/kg $1,290,000 $398,000 $201,000 $363,000 There is an additional, hidden performance advantage to a pumped engine, with increased thrust. Higher thrust

leads to shorter burn times to accelerate to the desired velocity. Over the long duration of a typical spacecraft burn, the vehicle traverses the curved path of a conic section, with thrust aligned with the desired direction of travel for only an instant. It was to minimize these turning losses that the Mercury Messenger spacecraft designers selected the Royal Ordnance 150 lbf Leros 1b engine over higher Isp, lower-thrust alternatives. Assessment of system benefit requires a trajectory simulation that is beyond the scope of this study, but it is comforting to know that there are additional system efficiencies out there to be realized.

V. Technology Development Required In order to realize the advantages outlined above, certain enabling technologies must be matured. They are:

A. Small Pumps for Storable Propellants There are a number of unknowns associated with small pumps for earth-storable propellants. The configuration

of the pump, centrifugal or positive displacement, is still subject to study and debate. Among positive displacement pumps, there are pistonless designs with some merits and variations on commercially available pumps that may be promising. Material compatibility with fuel and oxidizer has to be investigated in greater detail. A small piston-driven hydrazine pump has successfully operated on a small rocket14 but to date no NTO pumps have been made public. Electrically driven pumps must be demonstrated in a relevant environment and for any class of pumps, the efficiency and weight must be accurately characterized.

B. Start-up and Shut-Down Transients Pump-fed engines require more care in starting and shutting down than pressure fed engines. Convoluted

passages subject to heat soaking may be susceptible to thermal decomposition of hydrazine, which could lead to

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spontaneous energetic disassembly. Propellant lead and lag into the combustion chamber also affect the longevity of an engine. A progressive test program to demonstrate the start-up procedure is essential.

C. Power Conditioning Unit Aerojet has experience with design and qualification of Power Processing Units for electrical propulsion

systems, such as arcjets, ion engines and Hall current thrusters. Nevertheless development of this sort of assembly is never trivial.

D. High Temperature Chamber and Nozzle This study has assumed that increasing average chamber temperature, by decreasing FFC is a useful performance

improvement strategy. This requires that either existing materials be validated for use at still more elevated temperatures or that new materials or coatings be validated. Design improvements that will thermally isolate the chamber from the injector to minimize heat soakback are also needed. Better yet is the design and qualification of an appropriately sized regenerative chamber that has the potential to reduce chamber cost and simultaneously improve performance. The potential for cost reduction must be quantified to determine how much of the pump and power conditioning unit cost can be offset.

VI. Conclusion This study has shown that redesigning liquid, in-space propulsion engines to incorporate pumps has the potential

to deliver significant net performance improvements for optimized systems in general and for some specific vehicles in particular. In addition to the performance improvement resulting from higher specific impulse, inert component mass is substantially reduced. Higher Pc permits a similarly sized engine to yield increased thrust which can reduce turning losses and further increase system efficiency.

Higher Pc facilitates regenerative cooling, which has the potential not only to improve performance but to lower chamber cost to the point where adding pumps incurs little cost penalty.

Additional work is required to verify pump efficiency and weight, start-up and shut-down transients, power conditioning unit design, higher engine temperature engine operability, regenerative cooling efficiency and system reliability.

Acknowledgments The authors are grateful for the data and expertise regarding pumps for small engines provided by Ed Bennett,

Bill Buckley and Mark Nadolski, of Aerojet in Sacramento, California. The authors also wish to thank Carl Stechman for leading bipropellant engine development at Kaiser Marquardt

and bringing his expertise and mentoring spirit to Aerojet to raise up a new generation of engineers specializing in rocket propulsion.

References

1Daniel P. Thunnisen, Carl S. Guernsey, Raymond S. Baker, Robert N. Miyake, “Advanced Storable Propellants for Outer Planet Exploration” AIAA 2004-3488, 40th AIAA Joint Propulsion Conference, Fort Lauderdale, Florida 2J. Whitehead, “Reciprocating Pump Systems for Space Propulsion,” AIAA 2004-3836, 40th AIAA Joint Propulsion Conference, Fort Lauderdale, Florida 3Andrew Knight, “Designing and Testing a Lighter, Simpler, Less-Expensive Liquid Propellant Pump,” Journal of Propulsion and Power, Vol. 20, No 1, 2004, pp 141-154 4TDK’02, Two dimensional Kinetics (TDK) Nozzle Performance Computer Program, Software & engineering Associates, Inc., 1802 N. Carson Street, Suite 200, Carson City, NV 89701-1230. 5Dave Krismer, Antonio Dorantes, Scott Miller, Carl Stechman, Frank Lu, “Qualification Testing Of A High Performance Bipropellant Rocket Engine Using MON-3 And Hydrazine” 39th AIAA Joint Propulsion Conference, Huntsville, Alabama 6Takeshi Kanda, Goro Masuya, Yoshio Wakamatsu, Akio Kanmuri, Nobyo Chinzei, Masayuki Niino, “Effect of Regenerative Cooling on Rocket Engine Specific Impulse,” Journal of Propulsion and Power, Technical Notes, Vol. 10, No 2, March-April 1994, pp 286-288 7Kaiser Marquardt, “Mars Flyer Rocket Propulsion Risk Assessment,” NASA CR-2001-210710, April 2001.

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8Dieter K. Huzel, et al, Design of Liquid Propellant Rocket Engines, 2nd ed., National Aeronautics and Space Administration, Washington D.C., 1971, p 32. 9Phillip G. Hill, Carl R. Peterson, Mechanics and Thermodynamics of Propulsion, Addison-Wesley Publishing Company, Reading, Massachusetts, 1965, pp. 427-430. 10Carl Stechman, Horst Wichmann, “Economical High Performance Upper Stage Propulsion,” 38th Congress of the International Astronautical Federation, Brighton, Great Britain, October 10-17, 1987. 11M. B. Noel, “Experimental Investigation of Heat-Transfer Characteristics of Hydrazine and a Mixture of 90% Hydrazine and 10% Ethylenediamine,” NASA TR 32-109, Jet Propulsion Laboratory, June 27, 1961. 12Shuichi Ueda, Yukio Kuroda, Hiroshi Miyajima, “Bipropellant Performance of N2H4/MMH Mixed Fuel in a Regeneratively Cooled Engine,” Journal of Propulsion and Power, Vol. 10, No 5, Sept.-Oct 1994, pp 646-652. 13Messenger Launch Press Kit, August 2004, http://www.jhuapl.edu/newscenter/pressreleases/2004/MESSENGER_ Launch_Press_ Kit.pdf, 10 May 2006 14John C. Whitehead, Lee C. Pittenger, Nicholas J. Colella, “ASTRID Rocket Flight Test,” http://www.llnl.gov/etr/pdfs/07_94.2.pdf, 15 May 2006

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