Improvements to the Marketability of Hybrid Propulsion Technologies

Transcription

1 AIAA SPACE 2007 Conference & Exposition September 2007, Long Beach, California AIAA Improvements to the Marketability of Hybrid Propulsion Technologies Darren A. Kearney 1, Keith F. Joiner 2, Michael P. Gnau 3, and Mark A. Casemore 4 Lockheed Martin Space Systems Company Michoud Operations, New Orleans, LA, [Abstract] This paper will discuss technologies and approaches addressing improvements to hybrid propulsion platforms that will make these systems more affordable, practical and competitive for emerging boost applications. Specific topics to be addressed include improving overall mass fraction, system performance, packaging efficiencies, operations, and low cost/low rate production operations. Nomenclature L/D = the length of the rocket divided by the diameter of the rocket LO 2 = liquid oxygen LM = Lockheed Martin ISP = specific impulse (sec) lbm = pounds mass lbf = pounds force in 2 = square inches s = seconds HTPB = Hydroxyl Terminated Polybutadiene LMF900 = Lockheed Martin proprietary HTPB based fuel blend RR101 = Risk reduction test 1 (Falcon program) RR102+ = Risk reduction test 2 (Falcon program) HYSR = Hybrid Sounding Rocket (24 inch hybrid motor launched in 2003) HPDP = Hybrid Propulsion Development Program N 2 O = Nitrous oxide 250K = 250,000 lbf GO 2 = gaseous oxygen IRAD = independent research and development TVC = thrust vector control HTT = 10 inch hybrid motor FDN = forced deflection nozzle NDE = non-destructive evaluation GSE = ground support equipment I. Introduction YPICAL hybrid rockets are safer and have a higher average specific impulse (ISP) than solid rockets, and are T less complex and cheaper than liquid rockets, yet several specific obstacles need to be overcome to make them more marketable. Generally hybrid rockets are larger in volume than most solid rockets for the same overall 1 Aeronautical Engineer Sr, Advanced Programs, Lockheed Martin Space Systems Company, Michoud Operations, Dept 4160, Michoud Assembly Facility, P.O. Box 29304, New Orleans, LA Aeronautical Engineer Staff, Advanced Programs, Lockheed Martin Space Systems Company, Michoud Operations, Dept 4500, Michoud Assembly Facility, P.O. Box 29304, New Orleans, LA Acting Director, Advanced Programs, Lockheed Martin Space Systems Company, Michoud Operations, Dept 4500, Michoud Assembly Facility, P.O. Box 29304, New Orleans, LA Systems Engineer Sr, Advanced Programs, Lockheed Martin Space Systems Company, Michoud Operations, Dept 4200, Michoud Assembly Facility, P.O. Box 29304, New Orleans, LA Copyright 2007 by Lockheed Martin Space Systems Company - Michoud Operations. Published by the, Inc., with permission.

2 performance, due to the lower density of its propellants and the general requirement of more structure to separate its propellants and feed its fluid oxidizer into its fuel-lined combustion chamber. In addition due to hybrid s lower regression rates, past hybrids have been constrained in their volumetric shape where solids can range from very short, big diameter motors to very long, small diameter motors. Comparing length over diameters (L/D), hybrids are generally restrained to high L/D s which translates to long, skinny motors. Hybrid rockets are generally about the same volume of liquid motors, but again hybrids have typically been less capable of small L/D s where liquids can easily be tailored to fit most L/D requirements. Hybrid motors also generally have lower mass fractions than liquid motors due to the large amount of residual propellant that most hybrid rocket designs incorporate. Although hybrid motors have less fluids onboard than liquids, oftentimes hybrid motor feed systems are nearly as complex as liquid motors with multiple required fluids and pressurants to run the motors. With these issues along with the head start that solid motors and liquid motors have over the development and production experience of hybrid motors, hybrids are usually eliminated from new vehicle applications. Lockheed Martin has taken great strides in addressing these issues over the last few years, and many of these issues were directly tested and the risk reduced under the DARPA Falcon Phase II contract with Lockheed Martin. This paper will address the typical disadvantages, show how hybrid motors can be designed to overcome these disadvantages, and often show direct test data from Lockheed Martin research and development activities to confirm the feasibility of the designs listed. The paper will be broken down into sections based around the typical disadvantages or challenges of historical hybrid rockets, discussing volumetric envelope and packaging efficiencies, mass fractions and weight considerations, system performance and operations, and finally low cost / low rate production and operations. Most of these challenges must be overcome in order to make hybrid rockets more attractive to potential customers and improve the marketability of hybrid motors for market cases such as boosters, target drones, and second/upper stages. II. Volumetric Envelope and Packaging Efficiencies In order to evaluate the volumetric envelope of hybrid motors compared to solid and liquid rockets that are currently on the market, it is easiest to compare L/D s. For 1 st stage or booster rockets typical L/D s run between 4-12 for liquid and solid rockets as evidenced by the commercial vehicles public data shown in Table 1. Second stage motors and upper stage motors run between as seen in Table 1. All the motors in Table 1 have straight cylindrical geometry. Table 1. Length over Diameter for Liquid and Solid Motors 1 Liquid Rockets 1 st Stage (Core) Solid Rockets 1 st Stage or Booster Vehicle Stage L/D Vehicle Stage L/D Ariane 5 Core 5.6 Ariane 5 Booster 10.3 Atlas V 8.5 Athena (Castor 120) 3.8 Cyclone Atlas V Booster 12.8 Delta II 10.7 Delta II Booster 12.8 H-IIA 9.4 LM-2E Booster 6.8 K-1 (LAP) 3.0 M-V (Japan) 5.5 KOSMOS 9.24 Minotaur 4.5 LM-2C 7.66 Pegasus (Orion 50XL) 8.0 Scorpius (1 pod) 10.4 STS-SRB 12.0 Soyuz U Core ~10.8 Taurus(Orion 50SG) 6.8 Liquid Rockets 2 nd Stage (Upper) Solid Rockets 2 nd Stage (Upper) Vehicle Stage L/D Vehicle Stage L/D Ariane 5 Upper 0.8 Athena Upper 1.34 Atlas V (Centaur III) 4.2 Delta II (Star 48) 1.6 Cyclone 3 3 rd Stage 1.1 M-V (Japan) (2 nd ) 2.6 Delta II (2 nd Stage) 2.45 M-V (Japan) (3 rd ) 1.6 H-IIA (2 nd Stage) 2.3 Minotaur (SR-19) 3.0 K-1 (OV) 5.52 Minotaur (Orion 38) 1.4 Proton (3 rd Stage) 1.7 Taurus (3 rd Stage) 1.4 Soyuz U (3 rd Stage) 2.6 2

3 It becomes obvious that lower L/D s are more prevalent in the market and in order to compete against existing liquid and solid motors it would be best for hybrid motors to have L/D s less than By reducing the L/D of hybrids it might be possible to use more of the existing infrastructure from current liquids and solids, as well as reducing overall vehicle loads, easing transportation, and reducing the required control authority. An example of a hybrid motor in the public domain is the Hybrid Propulsion Development Program (HPDP) 250K hybrid motor 2, it was designed to have an L/D of approximately 20. The Hybrid Sounding Rocket (HYSR) sounding rocket as detailed in Ref. 3 had an L/D of In most cylindrical geometry hybrid motors the hybrid fuel grain is what dictates the basic outer diameter of the motor, and in the past it has been difficult to dramatically increase these diameters. Some motor designs like the SpaceShip One hybrid motor reduce the L/D of the hybrid rocket system by making the oxidizer tank much bigger in diameter than the hybrid motor. This was a successful design approach in that particular case due to the vehicle being shaped as an airframe, where the space around the motor could be filled with aero-control surfaces and such, but this uneven geometry is inefficient as a standalone rocket. The most marketable system would be to have the same diameter oxidizer tanks and hybrid motor (a straight cylindrical geometry) at the smallest L/D feasible. This section will detail various advances in the technology that help to bring the L/D of hybrid motors closer to those seen in existing liquid and solid motor systems. A. Fuel Grain Considerations Hybrid motors traditionally have lower regression rates than solid motors and this limits the diameter that a hybrid motor can achieve with a traditional wagon wheel motor design. With higher regression rates, a hybrid motor is able to be bigger in diameter and shorter in length for any given amount of fuel, which is preferred for most stages. Regression of the hybrid motor fuel is proportional to the mass flux of the motor at any time and at any axial location throughout the motor. Mass flux is defined as the rate of propellant flow across a cross sectional area. Typical regression rates are in the mils/s range, at flux levels between 0.15 and 0.30 lbm/in 2 s, where the simplest means to increase this regression is to increase the mass flux. In order to increase the flux this means reducing the cross sectional area of the exposed fuel surface or port, which often makes the rocket smaller in diameter. This change in flux generally does not alter the burn rate enough to make the total motor length any shorter. This technique of increasing the regression actually makes the L/D worse, if one assumes that the oxidizer tank and motor case are the same diameter. Lockheed Martin possesses a unique sizing tool, anchored in test data from over 750 tests ranging from 50 lbf to 250,000lbf thrust, which is able to quickly generate the basic dimensions and the fuel grain configuration of a hybrid rocket of a given impulse. A sample problem was put through this model to demonstrate how increasing the regression through increasing the flux actually makes the L/D worse. This sample problem is based around a 100,000 lbf thrust motor that burns for 60 seconds, where liquid oxygen (LO2) is the oxidizer and Hydroxyl Terminated Polybutadiene (HTPB) is the fuel. Some basic outputs from the model are in Table 2. The assumptions are that the LO2 tank is kept the same diameter as the hybrid motor, the hybrid motor keeps the same number of ports (9 total, 8 outer ports arranged around a center port in the traditional wagon wheel geometry), the LO2 tank and intertank volume remains the same between configurations, and the nozzle is the same. Table 2. Outputs from LM Proprietary Model Based on 100K 60 sec Hybrid Initial Flux Level Average Regression Rate Rocket Diameter Hybrid Motor Length Oxidizer Tank Length Total Rocket Length Void Space L/D lbm/in^2s in/s in in in in % % % % % % % % % % 25.6 Although the lower flux ranges lead to a lower L/D which is closer to contemporary liquids and solids, the cost of this is a much larger increase in void volume of the hybrid motor. Void volume is the amount of space within the 3

4 motor case that is not fuel, and in lower flux motors this means that the motor cases are mostly empty. Obviously the higher the void volume the less efficient the volumetric envelope of the motor is, as well as the more extensive and expensive the structure to contain this motor becomes. This also decreases the mass fraction. Advancements to increase the mass fraction of hybrid motors are discussed in the next section, and reducing the void volume is a fairly large contributor. It becomes obvious that the best hybrid motor would find someway to increase the flux and keep the void volume down, but at the same time allow for the maximum diameter so that the oxidizer tank can be larger in diameter, which both lead to smaller L/D s. In addition to increasing the flux to increase the regression rate, it is possible to increase the regression by the use of a faster regressing fuel such as paraffin. Lockheed Martin did several small scale tests as well as a fairly large scale LO2 test using paraffin as the fuel. The results from the test showed that paraffin was not as efficiently burning as HTPB, not as strong as LMF900 (LM proprietary HTPB blend with high strength), nor as storable and resistant to weather and temperature extremes as HTPB. In the Falcon program, as well as most military applications, high strength and resistance to temperature extremes are extremely important discriminators. Due to these facts, an alternate means to reduce the L/D of the motor using existing fuel formulations and regression rates was pursued. Lockheed Martin s final solution to maximize the volumetric envelope efficiency and achieve lower L/D s was to keep flux levels as traditional motors but add more rows of ports departing from the traditional wagon wheel design. Figure 1 shows a traditional wagon fuel grain versus a multi-row multi-port fuel grain design. Table 3 compares multi-row multi-port designs of the sample problem to a traditional wagon wheel with a flux level of 0.5 lbm/in 2 s. Traditional 1 Row 2 Row 3 Row Figure 1. Traditional Wagon Wheel Fuel Grain versus Multi-Row Multi-Port Fuel Grain Table 3. Traditional Hybrid versus Multi-Row Multi-Port Hybrid Outputs from 100K 60 sec Motor # Rows # Ports Rocket Diameter Hybrid Motor Length Oxidizer Tank Length Total Rocket Length Void Space L/D % % % % % 5.4 From Table 3, it is obvious that not only does adding additional rows of ports reduce the L/D, but that the void space in this particular example decreases as well. This void space decrease is not guaranteed in all cases, since it is a function of the regression rate of the fuel, the pre and post combustion chambers and any staged combustion systems that may be necessary. The decrease in L/D is guaranteed and decreases with each additional row added. At this point there must be a trade study to determine the optimum L/D based on fuel grain complexity, fuel strength, tooling costs, and several other factors related to hybrid motor design. Lockheed Martin performed a trade study and designed and successfully tested a 3 row 43 port fuel grain under the Falcon program for an estimated second stage rocket L/D of 5.8 with a total burn time of 220 seconds. Two risk reduction tests were performed to verify the model, test the fuel strength necessary for such a design, verify 4

5 proprietary LO2 injector design for even fuel grain burn, and verify combustion stability and high fuel efficiencies. Figure 2 shows a photo of the pre and post test condition of the 60 second RR101 test. The RR101 test was setup to be the same geometry as the final 220 sec duration fuel grain, and was run ¼ of its time or 60 seconds to ensure burn rates were as expected. Both risk reduction tests met all their objectives and successfully proved that such a multirow multi-port motor was practical and that LM would be able to predict the performance. RR101 Post-test (60 sec) Forward Aft Figure 2. RR101 Pre and Post Test Photos 3 Row 43 Port Fuel Grain B. Other Void Volume Reducing Considerations In addition to tailoring the fuel grain to make the overall rocket more volumetric efficient, Lockheed Martin performed a study to reduce the hybrid motors length and increase the packaging efficiency of the motor and intertank. LM liquid oxygen motors use a patented staged combustion to guarantee the combustion stability of the hybrid motors. Typical designs of this staged combustion system, as seen in the HPDP motor 2, have the heater motors located within the front dome of the hybrid motor as shown in Figure 3. This introduces several problems, such as higher fuel residual, added head end and injector complexity, and as the burn time increases lower heater capability. LM removed the heaters from the head end and introduced them into the fuel grain itself during the HYSR sounding rocket program 3. This reduced the overall motor residual, and head end complexity, but this introduced a different set of problems, such as increased drag due to oxidizer lines on the outside of the vehicle, and additional insulation within the motor itself. The Falcon program returned to a HPDP like configuration for its staged combustion design as it was felt to be the least risky option. After the RR101 test, another option was looked at to optimize the staged combustion system by placing the staged combustion motors within cans on the front of the hybrid motor. Figure 3 shows an example of this configuration. HPDP, Internal Heaters Can Heaters LO2 Inlet LO2 Inlet Heaters Heaters Potential Motor Case Reduction Figure 3. Regular Staged Combustion System versus Canned Staged Combustion System 5

6 The can approach reduced the overall length of the motor and decreased the fuel residual. It did cost extra structure, but the overall weight of the new system still showed a net positive weight benefit with the can staged combustion system. This decrease in motor case length was not carried into the intertank, because the intertank systems such as gas generator and turbopump system could be interspaced between the protruding heaters. This allowed for a lower void volume within the intertank, and was considered a benefit and good option for reducing the overall L/D. The final benefit to the can approach was to alleviate some of the heater motor performance loss associated with longer burn times. When the staged combustion motors exhaust straight into the head end they become less effective as the heater motor ports grow larger. This is due to the reduced velocity of incoming hot exhaust products versus the constant oxidizer velocity. The exhaust products are less capable of penetrating the oxidizer stream and thus mixture of the two products is reduced. The can approach was to have a constant area duct that the exhaust products would exit such that the velocity would not be as dramatically reduced. Although the can approach traded out as a much better solution than the standard staged combustion approach and reduced the L/D and volumetric efficiency it was not chosen for the Falcon motor as a risk savings method. Although some test data was available on this approach it was decided to proceed with the standard approach to avoid additional confirmation testing. Future LM hybrid designs will continue to evaluate this trade with the design s specific requirements as a potential volumetric efficiency and performance increasing measure. A final way to reduce the L/D of the rocket, with the added benefit of increasing the mass fraction, is to use a pump-fed system versus a pressure fed system. This does not apply to self-pressurizing propellants such as N 2 O (nitrous oxide). For the pressure fed system, an additional high pressure vessel is required in order to pressurize the oxidizer tank, and this is often significant enough to add a fair amount to the length of the rocket. Figure 4 shows a pressure-fed vehicle versus a pump-fed vehicle with identical performance. Hybrid Motor Oxidizer Tank Pump Fed L/D = 6 Hybrid Motor Oxidizer Tank Pressurant Tank Pressure Fed L/D = 7 Figure 4. Pressure Fed versus Pump Fed Vehicle of Same Performance It is obvious that using a pump fed system reduces the length of the vehicle. The HPDP 250K program recognized this advantage and the 250K motor was designed to incorporate a turbopump to feed the LO 2 into the motor. The turbopump was to be driven by a hybrid based steam generator in order to maintain the basic safety and inertness of the overall rocket. There was a disadvantage to this, in that a water storage vessel was going to be required and this made the intertank a little larger to accommodate the water storage vessel s volume. In addition the HPDP motor was required to have GO 2 storage vessels onboard in order to run the staged combustion system. To further reduce the volume of the turbopump and staged combustion systems, Lockheed Martin pursued a new pump-fed design for the Falcon program that was highly integrated. It used a hybrid motor to heat up LO 2 from the main oxidizer tank and then passed this hot GO 2 into a turbine to run the turbopump of the motor, as well as using this GO2 for tank pressurization and staged combustion oxidizer. This system significantly increased the volumetric efficiency of the rocket by removing pressurant vessels for oxidizer pressurization, as well as pressurant vessels for the staged combustion system. This highly integrated system was able to fit within a fairly compact intertank. Testing was performed on various aspects of this system under LM IRAD, with the most significant testing, the hot GO2 generator, successfully tested three times. In order for hybrid motors to be more marketable, it is a good idea 6

7 to move to pump fed systems (at least for large stages) and to integrate the oxidizer feed systems to reduce additional liquids and their associated storage vessels. Additional information on this system will be described in the system performance section. Lockheed Martin has significantly increased the state of the art of hybrid technology to increase the volumetric efficiencies of hybrid motors. Through the research and successful testing performed under the Falcon program, hybrid motors were designed to be comparable to the L/D s normally associated with solid and liquid rockets. Many of the previously listed changes also make hybrid rockets more comparable to solids and liquids in other performance criteria. The next section of this paper will focus on the advances that have been made on making hybrid rocket mass fractions more competitive. III. Mass Fraction and Weight Considerations Similar to the volumetric efficiency it is a good idea to compare hybrid motors to existing liquid and solid rocket systems when looking at mass fractions. Typical solid motors have fairly good mass factions due to minimum structure requirements. The structure requirements of most solids is directly correlated to the fairly high density that solid fuel has, as well as the fact that all the oxidizer and fuel is mixed together and as such does not need to be separated with additional structure. The main limiter to mass fraction on solids is due to the high pressures that the motor cases must withstand. Liquid motors have good mass fractions mostly due to the low pressure vessels that contain the liquid oxygen and fuel (on turbopump systems). The combustion chamber is the only vessel that must be at a high pressure and this is generally fairly small compared to the overall remaining structural elements. Liquid engines are limited in that most of the propellants are at fairly low densities and require larger volume and more structure. Table 4 shows the mass fraction of a selection of solid and liquid motors currently on the market. Table 4. Mass Fractions of a Selection of Liquid and Solid Rockets 1 Solid Motors Liquid Motors Vehicle Stage Mass Fraction Vehicle Stage Mass Fraction Ariane 5E Booster 0.86 Ariane 5E Core 0.91 Athena Stage 1 (Castor 120) 0.92 Ariane 5 ESC-B 0.88 Delta II Booster (Gem 40) 0.90 Atlas V Core 0.93 H-IIA SRB-A 0.86 Delta II Core 0.94 LM-2E Booster 0.92 Delta II 2 nd Stage 0.86 M-V Stage H-IIA Stage M-V Stage 0.91 K-1 LAP 0.91 Minotaur Stage 1 (M-55A) 0.90 K-1 OV 0.90 Minotaur Stage 2 (SR-19) 0.89 KOSMOS Stage Pegasus Stage 2 (Orion 50XL) 0.90 Proton Stage Pegasus Stage 3 (Orion 38) 0.88 Scorpius Stage 1 (6 pods) 0.85 STS SRB 0.85 Scorpius Stage From Table 4, it is obvious that for hybrids to be competitive with liquids and solids they should have mass fractions as close to the 0.84 to 0.92 range as is affordably feasible. In the case of the HPDP 250K booster the mass fraction was expected to be approximately 0.8. The HYSR sounding rocket had a mass fraction of Under the Falcon program, the first stage mass fraction was going to be approximately 0.89 while the second stage mass fraction was going to be approximately Part of this dramatic increase in hybrid mass fractions was due to the techniques mentioned earlier. Multi-row multi-port fuel grains and highly integrated oxidizer feed systems reduced the structure requirements, thus increasing the mass fraction. In addition, the integrated feed systems removed additional weight by eliminating heavy pressurant vessels. This section will focus on additional techniques and design considerations that helped Lockheed Martin make the mass fraction of its Falcon hybrid motors much more competitive. A. Fuel Residual Considerations Hybrid motors have historically had fairly large residuals which directly correlate to mass fraction. Mass fraction is based on the initial vehicle weight and the burn out weight, and residual fuel is part of the burn out weight 7

8 of the vehicle. Lockheed Martin developed a technique to reduce the fuel residual of hybrid motors through the use of planned fuel expulsion, with details of this process outlined in Ref. 4. The total fuel grain fuel weight is allowed to be much smaller than a typical hybrid motor because neither structural stiffeners within the fuel grain or additional fuel between ports are required to maintain the structural integrity of the fuel grain. Structural integrity of the fuel grain is important, because in previous hybrid tests the fuel could chunk fairly large pieces of the hybrid fuel grain out of the nozzle. These chunks would increase the chamber pressure dramatically for a short period of time (possibly rupturing the motor case unless it was designed for a higher pressure than nominal performance), and with the fuel webs no longer in the motor the overall performance of the motor would be dramatically reduced for the remaining burn time. The planned fuel expulsion technique uses two significant improvements in hybrid technology to maintain the structural integrity and performance of the fuel as long as possible, at which point very small slivers of unused fuel are ejected through the nozzle. The performance of the rocket takes into account this eventual small loss in performance and evaluates the slivers to ensure that the chamber pressure is not significantly increased. The first improvement that makes planned fuel expulsion feasible was to increase the strength and elasticity of the fuel. Lockheed Martin used a design of experiments to determine the strongest HTPB based fuel that could be cast while still maintaining the same regression characteristics of the regular HTPB that LM has over 800+ tests experience with. The final fuel formulation, designated LMF-900, increase the tensile strength of the fuel to 900 psi, which was a 600% increase over existing room temperature cured HTPB and increase the elasticity of the fuel by 650%. As detailed in Ref. 4, a stronger, more elastic fuel allows for the webs between fuel ports to burn to much smaller thicknesses before final expulsion from the motor. In previous motors, chunks much greater than one inch in thickness were routinely seen exiting the motor unburned, while with the stronger fuel the thickness was less than 1/8 thick. These small slivers were shown in Ref.4 to have no noticeable effect on vehicle chamber pressure, and the small change in performance was able to be modeled accurately. Figure 5 shows the strength and elasticity of the new LMF-900 fuel over previous LM fuel blends and the HPDP fuel blend Average sample from verification matrix tested at 20 in/min Stress, psi Improved Hybrid Fuel Formulation HPDP Ultimate The Improved Hybrid Fuel Formulation Yields a 600% Increase in Tensile Strength and a HTPB Fuel 650% Increase in Strain Sample Strain, in/in Figure 5. Tensile Strength and Strain of LMF-900 Fuel vs. Previous Fuels The second improvement that makes planned fuel expulsion feasible was the use of a stress model to accurately predict the failure point of the fuel grain. Ref. 4 shows how the stress model was created and how to incorporate the 8

9 stress model into the fuel grain ballistic model to accurately predict the change in performance when the fuel slivers. Ref. 4 shows test data and photos from a small scale two port hybrid motor, but LM used this same procedure to determine the planned fuel expulsion of the Falcon motor and to dramatically reduce the residual of its fuel grains. The Falcon motor was to reduce the fuel grain residual from typical hybrid fuel grain residuals of 12+% to a fuel grain residual of approximately 3%. This planned fuel expulsion was tested in the 120 second RR102+ test under the Falcon program and was completely successful. It was found through the use of the stress model on the multi-row multi-port design that the only way to burn the fuel grain and maintain structural integrity of the grain was to burn the motor preferentially from the center row out. This would ensure that the minimal amount of fuel would be slivered, that the motor case would stay intact and the overall performance would be maximized. In order to ensure that the burning was preferential some new LO2 injector design work was also performed. The combination of a new LO2 injector design and the spacing of the ports allowed the fuel grain to burn from the center out towards the motor case. Figure 6 shows the post test pictures of the RR102+ motor as well as a normalized pressure trace to prove that the planned fuel expulsion technique is successful at even the most complex fuel grain geometry. This technique along with a strong fuel is very effective at reducing the fuel grain residual of a hybrid fuel grain directly leading to a higher mass fraction over historical hybrid rockets. Figure 6. RR102+ Post Test Photos and Chamber Pressure 9

10 A final approach that is used in LM hybrid motors to reduce the residual of the hybrid motor is the use of high temperature insulation within the motor case. A simple layer of silica phenolic or other insulator can be placed on the motor case mandrel before wrapping the composite for the motor case. This insulation allows the fuel grain to be burned all the way to the walls of the motor case. Insulation is necessary in most hybrid motors, due to the fact that hybrids generally burn more at the back than at the front. Using an insulator is lighter because otherwise a significant thickness of fuel would be required at the back of the motor to ensure that the motor case does not have a burn-through. The 1/8 to ¼ thick insulation is much thinner and lighter than the often 3/4 to 1 of fuel that would be required to ensure motor safety. B. Structural Considerations Another advance in hybrid motor manufacturing has led to an increase in mass fraction. In most hybrid designs, the motor case is separated into 3 distinct sections, the head end or pre-combustion chamber, the center section or fuel grain, and the aft end or post combustion chamber. These three elements are then generally bolted together after casting to finish the motor case. The production process is usually performed in this method because the pre-combustion and post combustion chambers are much larger than the fuel grain individual ports so that if they were cast in one pour, the tooling for the chambers would be stuck in the motor. The need for two joints in the past leads to an increase in overall weight. LM has pursued a soluble tooling approach for the pre-combustion chamber so that the fuel grain and head end can be cast at the same time and thus reduce at least one joint from the motor case. One problem that had to be overcome was finding a water soluble tooling material or a coating for the tooling that would not outgas during the motor casting process and introduce voids into the hybrid fuel. In addition the material or the coating for the soluble material had to not soak up fuel during casting and form a rind that would not be removed when the tooling was dissolved. The material also had to be able to withstand the curing temperatures that LMF 900 is subjected. Finally the material had to be easily dissolved with water for ease of manufacture. LM spent the greater part of 2006 running experiments on soluble tooling and went through various off the shelf solutions before finally choosing a proprietary material blend from one of our motor case vendors. The sand based tooling material is easily manufactured at nearly any size, able to stand up under its own weight, and washes out easily with water. The constituents of the material are such that there is no need for special disposal methods while it is being dissolved. In order to ensure the fuel does not have voids introduced into it, a LM proprietary coating is applied to the finished tooling mold prior to casting. Through the use of water soluble tooling, LM has been able to remove the head end-fuel grain joint and decrease the overall structure weight of the motor case. In the Falcon program this change in the manufacturing of the fuel grain and motor case led to a 1% decrease in the second stage motor weight. With mass fraction such a critical component in most high performing rockets, a 1% decrease in overall stage weight is an attractive option for a minimal change in the manufacturing flow. Another technology that LM evaluated to increase the mass fraction was the use of composite tanks for the liquid oxygen and hybrid motor. Hybrid motor cases have been made out of composite in many previous rockets and this standard approach was baselined for the Falcon program. Filament-winding technology is well established and weight reductions between 20-40% over aluminum motor cases can be expected. LM has also developed and tested cryogenic composite liquid oxygen tanks. These tanks were demonstrated under the X-34 project with over 50 cryogenic cycles with LO2 and this composite expertise can be applied to increase the overall mass fraction of hybrid rockets. In order to take advantage of the 20-40% weight reduction available for LO 2 tanks on hybrid motors, two primary issues had to be overcome. These issues are cryogen containment and chemical compatibility. Cryogen containment is maintained by reducing leak paths through the control of defects and microcracking in the composite. Compatibility is assured through the use of proper chemical compatibility requirements and testing the composite through standard mechanical impact, friction, pyrotechnic shock, and high velocity particle impact tests. Lockheed Martin has over twenty-five years of continuous development on composite structures, and details of this capability and the solutions to overcome the containment and compatibility problems with LO 2 cryogenic tankage can be found in Ref. 5. LM has increased the mass fraction of hybrid rockets to make them more competitive with liquid and solid systems and some of these advances were demonstrated during the Falcon program. The planned fuel expulsion technique coupled with a much stronger fuel were proven to dramatically reduce residual on hybrid rockets. Soluble tooling approaches were developed under IRAD to decrease motor case weight through the elimination of joints, and our historical success with composite oxygen tanks also increased the mass fraction of hybrid rockets. 10

11 IV. System Performance and Operations In this section there will not be a direct correlation or comparison to liquid or solid rockets, but instead this section will concentrate on system improvements that make hybrids more attractive. LM has spent the last 15 years working on improvements to make hybrids more stable and efficient, and the first part of this section will discuss these efforts and how they lead to a more attractive hybrid rocket solution. The second part of this section will discuss how under the Falcon program, LM improved oxidizer feed systems, nozzle designs, and tailored the fluid system requirements to make hybrid rockets achieve higher performance. A. Motor Stability and Efficiency Considerations LM developed analytical techniques and a comprehensive analysis approach to evaluate the performance of hybrid motors, and Ref. 6 outlines this approach. One of the first criteria that a rocket must meet is a stable combustion environment. This is necessary for predictable rocket performance and lower weight structures (as described in the previous section) Lockheed has worked on the stability issue for over a decade with one of the largest successes documented in the HPDP program. Under the Motor 2 Design Basis and History in Ref. 2, there is a large section describing the methodology and science behind the LM patented staged combustion motor concept. The results from the three staged combustion HPDP tests presented interesting results in terms of stability and efficiency. Figure 7 shows the sea level thrust traces from all of the HPDP 250K tests. Motor 1 (passive head end) and Motor 2 (staged combustion head end) are on the same chart so that the differences in performance between a LM designed motor and a typical industry motor can be compared. The second test of HPDP Motor 2 had stability issues that were expected by the time of the test. The HPDP motors had to be designed and fabricated years in advance of the test, and the staged combustion approach was in its infancy at the time of design. Through other tests held between design and test, LM developed a minimum heat requirement for stable combustion. HPDP Motor 2 Test 2 was expected to be unstable because the previously designed heater motors were determined to only provide enough heat for 20 seconds of stable burn. Approximately 8 seconds into test 2 (23 seconds of burn on the same fuel grain), the motor went dramatically unstable. Approximately a year later, additional funding was acquired for another test, and LM was given the chance to prove that it had since worked out the stability requirement. LM refurbished the head end of the motor with a heater design that was expected to maintain stability for the remaining burn time. The motor was then tested, and from Figure 7 HPDP Motor 2 Test 3, it is obvious that the motor ran extremely stable. LM has used the proven successful staged combustion motor process since that time with exceptional stability proven on HYSR, and Falcon programs. 350 K 400 K 250 K Sea Level Thrust (lbf) 300 K 350 K 250 K 200 K 150 K Motor 1 Motor 2 Sea Level Thrust (lbf) 150 K 100 K 100 K 50 K 50 K 0 K K Time (s) Time (s) 300 K 250 K 200 K Sea Level Thrust (lbf) 200 K 150 K 100 K 50 K 0 K Time (s) HPDP Motor HPDP Motor 2 Test 2 HPDP Motor 2 Test 3 Figure 7. HPDP Motor 2 Test Data Stable combustion has also been proven to increase the efficiency of hybrid motors with a listing of ISP and Cstar efficiencies listed in Table 5. High efficiency hybrid motors are desired because it reduces the overall weight of the rocket for a given payload. Table 5. Staged Combustion Motor Data Test Series Fuel Combination # Tests Geometry ISP eff C* eff HPDP 11 HTPB GO2/LO2 60 Single port 95% 99% HPDP 24 HTPB LO2 56 Multi-port 95% 98% HPDP 250K HTPB LO2 4 Multi-port 94% 98% HMF HTPB /AL LO2 16 Single port 95% 98% HYSR HTPB/AL LO2 6 Multi-port 95% 98% RR10x LMF900 LO2 2 Multi-row Multi-port 97% 99% 11

12 B. Hybrid Improvements Developed Under Falcon Program The hybrid platform offered a chance to change the paradigm of rocket system performance and operations under the Falcon program. The Falcon program required a fairly cheap rocket to launch within 24 hours of call up from within the continental U.S. In order to meet these strict requirements, LM thought that the minimization of fluids required would significantly decrease the vehicle loading times and reduce the ground support equipment (GSE) that would be required to support the launch. LM decided to pursue a single fluid approach. LO2 was required as the oxidizer for the rocket due to its high ISP and relatively cheap production and handling costs and easy availability. The hybrid feed system was then designed to function on the single fluid, LO2, with a single vessel, the main LO2 tank, required to hold the fluid. This was accomplished through the use of a revolutionary change in how the turbopump was driven. The turbine would be designed to run on a hot oxidizer instead of the more ubiquitous hot fuel, with a hybrid motor used to heat the liquid oxidizer from the LO2 main tank. In addition, that hot GO2 would then be used to feed the staged combustion motors, thus reducing the structure and fill needs for separate GO2 tanks. A simplified, partial schematic of how this system works is shown in Figure 8. GO2 For Tank Pres GO2 LO2 Pump GO2 From Subsystem Hybrid Fuel Hybrid Fuel GO2 Overboard GO2 Turbine LO2 GO2 LO2 Hybrid Motor Mixing Chamber GO2 For Staged Combustion Figure 8. Simplified Schematic of Hot GO2 Generator Driven Turbopump System This system simplified the requirements on the turbopump, because it no longer had to seal a hot fuel on the turbine from an oxidizer on the pump side of the turbopump. Since the fluid was the same on both components of the turbopump, a simpler seal design was created. The riskiest component of this system was identified to be the hybrid gas generator, and this component was tested under LM IRAD funding during the Falcon program. In order to keep the LO2 from immediately igniting when introduced to the hybrid exhaust, the hybrid motor was run oxidizer rich, and the mixing chamber was made from Monel so that the container itself would not burn. The final temperature of the oxygen after its test was as predicted and expected. Figure 9 shows a cross section of the tested hot gas generator, and a photo of one of the tests that was performed. Further detail of the gas generator can be found in Ref

13 GO2 Inlet Hybrid Fuel Grain Insulation LO2 Inlet Mixing Chamber Exhaust Figure 9. Gas Generator Cross Section and Test Photo With the new single fluid system for pump fed motors, hybrids are becoming more competitive, reducing GSE requirements, cost of turbo-machinery, and weight of the vehicle through reduction of extra structure. Another system that was developed during Falcon to make hybrids more competitive includes the use of Aerojet s hot gas injected thrust vector control system with a forced deflection nozzle. This kind of thrust vector control (TVC) system is a more effective system for TVC over regular liquid injected or gas injected TVC systems because it uses a much hotter gas (straight out of the hybrid combustion chamber) to divert the main flow in the nozzle. Aerojet has had good success with this system on solid rocket motors but they never tested the system in a hybrid rocket environment. Under the Falcon program, LM worked toward identifying the materials for a hot gas valve that would stand up to the harsher hybrid motor environment. The hybrid motor environment is approximately 1000 degrees Fahrenheit hotter and contains more free oxygen. In order to identify this material, a small material test plan was drawn up and completed under IRAD funding. The tests were completed using HTT 8 hardware and the materials tested were 4D carbon-carbon, graphite, and rhenium. Rhenium was chosen as the final material for the pintle and seat within the hot gas valves. 4D carbon-carbon was chosen as the throat material for the nozzles. LM intends to test a hot gas valve this year to see how it performs in a hybrid rocket environment. The forced deflection nozzle was chosen for the Falcon stages because it simplified TVC, was integrated straight into the hybrid motor casing (which eliminated weight), it provided altitude compensating capabilities up to the maximum diameter of the nozzle, and was determined to be significantly less costly than conventional bell nozzle configurations with TVC. The increase in performance from such a nozzle was worth the risk reduction testing that needs to be done in order to ensure it functions well in a hybrid motor. Through the choices of more compact, combined systems and new technologies hybrid motors are much more competitive in terms of performance with solid and liquid motors. The systems outlined in this section also do not increase the cost of the vehicle significantly, which leads to a more marketable hybrid rocket. V. Low Cost / Low Rate Production Operations One of the biggest advantages of hybrids over liquids and solid rockets is the simplicity and safeness at which they can be produced. This section will discuss these advantages as well as outline how some advances have been made to make even low rate production efforts relatively low cost. This makes hybrid rocket especially attractive to potential customers. Perhaps the most obvious advantage of hybrid motors is the simplicity at which they can be cast. It is literally as easy as mixing up a batch of fuel, pouring in into a mold under vacuum, and curing the fuel. Solid motors have very strict mixing requirements and safety considerations so that the mixed fuel and oxidizer are carefully casted to avoid voids, cracks, debonds, and possible detonation and explosion of the fuel during mixing. Most solid rocket plants are placed in remote locations, either in the desert or at the top of mountains and generally must produce fuel in big volume to be cost effective. The hybrid motors LM casts are designed, cast, and assembled all in a shirt sleeve area 13

14 without any major infrastructure requirements. In the past, LM hybrid motors were small enough that a few pours from a 50 gallon mixer were sufficient to cast the motors. The risk reduction motors for Falcon were significantly bigger and required a low cost, high volume mixing capability. The simplest, cost effective means to do this was the use of a continuous, static mixing system. Figure 10 shows a photo of the simple system that uses standard off the shelf industrial tanks and pumps, and is capable of casting any size hybrid motor. Additional tanks can be added of the two fuel components to meet any requirement. Tank A is comprised of the HTPB fuel, crosslinker, and carbon opacifier, while Tank B is comprised of the isocyanate and antioxidant. The small mixer is used to pre-mix the part A&B components before filling the storage tanks. The fuel components in each tank can sit for weeks before casting because the active components of the fuel do not come together until it is time to pump them through the static mixer and into the hybrid rocket mold. This simple system is very attractive because it allows production of large or small hybrid rockets at any production rate. Tank A 55 Gallon Mixer Static Mixer Tank B Test Article Vacuum Pumps Figure 10. LM Hybrid Rocket Casting Setup In addition to low cost casting of hybrid motors, the non-destructive evaluation techniques on hybrid rockets do not have to be as robust as a solid rocket motor. The main reason for this is that voids in the volume of the hybrid are typically not huge contributors to performance variations unless they are located close to the interior motor case wall. These voids are easy to spot with a cursory observation technique such as digital X-ray. Hybrid motors do not need expensive NDE techniques such as CT scans, or magnetic resonance imaging. Another device to ensure good hybrid motors is process control, such as ensuring that the motor case is cast under vacuum, and that all the fuel components are within their shelf life and free of additional moisture. To further reduce the cost of NDE, and increase the probability of success, high temperature insulation has been developed and tested at LM that reduces the likelihood of motor case burn-throughs. This insulation has already been outlined in the section above about lower fuel residual. Between the lower critical void size requirement and having protective insulation, NDE is much easier for hybrid rockets. Other factors that reduce the cost and complexity of the hybrid rocket system, such as dissolvable tooling, and single fluid systems, have been outlined for the performance improvements they make to hybrid rockets. These two factors are also cheaper and easier to produce in lower volumes, and should not be looked over for their contributions. The final contribution to cost and production improvements over past rockets is the use of horizontal integration for stage and payload connections. Horizontal integration is an obvious technique to reduce the costs of launching a rocket due to the smaller infrastructure requirements and most rocket systems are trying to incorporate horizontal integration in their current designs. 14

15 VI. Conclusion Hybrid rockets have in the past not been viewed as competitive, or as well developed, and therefore not as marketable as solid or liquid rocket systems. This paper discussed how Lockheed Martin has addressed the many issues that have been associated with hybrid rockets and how the problems have been solved with data to support the claims. It is LM s belief that hybrid rockets are viable rocket systems that can greatly reduce the cost and complexity over today s current systems and the marketability of hybrid rockets is greatly improved. Through LM s efforts, under IRAD development and the Falcon program, many new techniques were developed and can be applied to make hybrid motors comparable with other systems. Through successes in the commercial suborbital tourism industry, interest in hybrid motors has grown stronger, and with the outlined improvements developed by Lockheed Martin, it is time to capitalize on hybrid motor technology. D. Kearney thanks: Bobby Biggs Joseph Arves Acknowledgments K. Joiner thanks: Tim Knowles Mike Gnau thanks: Bob Sackheim DARPA and the Air Force for support in advancing hybrid technology References 1 Isakowitz, S. J., Hopkins, J.B., and Hopkins Jr., J. P., International Reference Guide to Space Launch Systems, 4 th ed., AIAA, Reston, Virginia, Story, G., Zoladz, T., Arves, J., Kearney, D., Abel, T., Park, O., Hybrid Propulsion Demonstration Program 250K Hybrid Motor, 39 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, AIAA , Huntsville, AL, Arves, J. P., Gnau, M., Joiner, K., Kearney, D., McNeal, C., Murbach, M., Overview of the Hybrid Sounding Rocket (HYSR) Project, 39 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, AIAA , Huntsville, AL, Kearney, D., Geiman, W., Accounting for Planned Fuel Expulsion by Hybrid Rockets, 41 st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, AIAA , Tucson, AZ, Achary, D., Biggs, R., Bouvier, C., McBain, M., Lee, W., Composite Development & Applications for Cryogenic Tankage, 2005 National Space & Missiles Materials Symposium (NSMMS), Las Vegas, NV, June 26-July 1, Arves, J., Jones, H., A Standardized Technique for Evaluating Hybrid Rocket Motor Performance, 33 rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Seattle, WA, July 6-July 9, Dupuis, T. Knowles, T., Oxygen Rich Hybrid Gas Generator, 42 nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Sacramento, CA, July, Knowles, T., Kearney, D., Overview of 10 inch Diameter HTPB Hybrid Motor Testing with Liquid Oxygen at Stennis Space Center, 41 st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, AIAA , Tucson, AZ,