Concept Study of an ARES Hybrid-OS Launch System

Concept Study of an ARES Hybrid-OS Launch System Jon G. Wallace 1, John Bradford 2, A.C. Charania 3, and William J.D. Escher 4 SpaceWorks Engineering, Inc. (SEI), Atlanta, GA, 3338 Dean Eklund 5 Aerospace Propulsion Division, Air Force Research Laboratory, Wright-Patterson AFB, OH, 45433 The Affordable Responsive Spacelift (ARES) program is part of a broad initiative by the United States Air Force to improve the responsiveness, flexibility, and affordability of space access capabilities. Based on initial studies, including Analysis of Alternatives (AoA) activities, the Air Force has concluded that the ARES program should pursue development of a hybrid launch system which utilizes a reusable booster and an expendable upper stage. The design philosophy behind this configuration contends that significant cost savings can be achieved compared with an all-expendable approach by reusing two thirds of the hardware mass of the launch system after each flight. In addition, substantial benefits to operational cost and complexity can be realized if the trajectory of the reusable booster is tailored such that a complicated thermal protection system is not required. SpaceWorks Engineering, Inc. (SEI) has supported this effort through a recent design study of an ARES Hybrid Operational System (OS) launch architecture. This study included vehicle performance, life cycle cost, operations, safety, and reliability analysis. Starting with the set of baseline groundrules and assumptions, SEI proceeded to define a feasible upper stage system, design the reusable booster airframe, analyze the main propulsion system for the reusable booster, and determine a means of returning the booster to the launch site after separation from the upper stage. Based on the results of this conceptual design activity, SEI has been able to answer several key questions regarding the feasibility and viability of an ARES Hybrid-OS concept. For instance, the trajectory simulations and aeroheating analysis performed as part of the design closure reveal that it would be possible to forego an extensive thermal protection system on the booster one of the linchpins of the highly operable ARES approach. SEI s analysis also explored the potential synergy between low cost commercial small payload launch vehicles and the design requirements for the Hybrid system s expendable upper stage. In addition, a detailed examination of the ARES Hybrid-OS booster main propulsion elements enabled SEI to refine previously vague design guidelines for development of these liquid locket engines. Nomenclature ACS = attitude control system AoA = Analysis of Alternatives; angle-of-attack ARES = Affordable Responsive Spacelift CONOPS = concept of operations CONUS = continental United States EHA = electro-hydraulic actuator EMA = electro-mechanical actuator EPC&D = electrical power conditioning and distribution ETO = Earth-to-orbit 1 Senior Project Engineer, Advanced Concepts Group, 12 Ashwood Pkwy., Suite 56 and Member AIAA. 2 President, 12 Ashwood Pkwy., Suite 56 and Senior Member AIAA. 3 Senior Futurist, Economic Engineering Group, 12 Ashwood Pkwy., Suite 56 and Senior Member AIAA. 4 Senior Technical Fellow, 12 Ashwood Pkwy., Suite 56 and Senior Member AIAA. 5 Aerospace Engineer, Propulsion Technology Branch, AFRL-WPAFB and Senior Member AIAA. 1

GLOW = gross lift-off weight GOX = gaseous oxygen IOC = initial operating capability LOX = liquid oxygen MECO = main engine cut-off MER = mass estimating relationship OS = operational system RP-1 = Rocket Propellant 1 RTLS = return to launch site SLS = sea level static TVC = thrust vector control VTHL = vertical take-off, horizontal landing I. Introduction HE Affordable Responsive Spacelift (ARES) program is part of a broad initiative by the United States Air Force T to improve the responsiveness, flexibility, and affordability of space access capabilities. Termed Operationally Responsive Space (ORS), this policy initiative has led to several forward-looking programs including the joint Air Force DARPA Falcon program and the Affordable Responsive Spacelift (ARES) endeavor. 1 Specifically, the ARES program calls for the development of a new medium payload Earth-to-Orbit (ETO) launch system that leverages the advantages of both reusability and expendability. Whereas FALCON is intended to provide low-cost space access for small payloads (<1, lbs), ARES seeks to bring about a similar transformation in the Air Force s medium payload (~15, lbs) capability to low Earth orbit. Based on initial studies, including Analysis of Alternatives (AoA) activities, the Air Force has concluded that the ARES program should pursue development of a hybrid launch system which utilizes a reusable booster and an expendable upperstage. The design philosophy behind this configuration contends that significant cost savings can be achieved compared with an all-expendable approach by reusing two thirds of the hardware mass of the launch system after each flight. In addition, substantial benefits to operational cost and complexity can be realized if the trajectory of the reusable booster is tailored such that a complicated thermal protection system is not required. SpaceWorks Engineering, Inc. (SEI) has supported this effort through a recent design study of an ARES Hybrid Operational System (OS) launch architecture (see Fig. 1). This study included vehicle performance, life cycle cost, operations, safety, and reliability analysis. Starting with the set of baseline groundrules and assumptions, SEI proceeded to define a feasible upper stage system, design the reusable booster airframe, analyze the main propulsion system for the reusable booster, and determine a means of returning the booster to the launch site after separation from the upper stage. Figure 1. SEI ARES Hybrid-OS launch system concept. 2

II. ARES Hybrid-OS The SEI ARES Hybrid Operational System (Hybrid-OS) launch architecture is a multi-stage, partially reusable, all-rocket launcher. SEI s design approach is centered on the synergistic use of both new and existing technologies in the context of an innovative launch system with the objective of achieving high levels of responsiveness, affordability, and operability. Based on the technology assumptions, an initial operational capability (IOC) for this system in the year 211 to 215 is envisioned. A. Key Technologies The SEI ARES Hybrid-OS concept is enabled through a combination of new and existing technology solutions in the areas of propulsion, materials, structures, avionics, and integrated vehicle health monitoring (IVHM). Key design features and technologies for the booster include: Vertical take-off, horizontal landing (VTHL) All-new, domestic (United States) LOX/RP-1 main rocket engines Aluminum primary and secondary airframe structure Integral, cylindrical Aluminum fuel and oxidizer tanks Two (2) low-bypass turbofan engines for booster flyback (e.g. F118-GE-1) EHA s (electro-hydraulic actuators) for control surfaces Integrated Vehicle Health Monitoring (IVHM) systems Key features and technologies for the upperstage system include: All-rocket, two stage upperstage derived from Space Exploration Technologies (SpaceX) Falcon-1 launch vehicle SpaceX Merlin LOX/RP-1 engines on 2nd stage SpaceX Kestrel LOX/RP-1 engine on 3rd stage EMAs (electro-mechanical actuators) for pitch and yaw control B. Mission Overview The mission of the ARES Hybrid-OS concept is to deliver 15, lb of payload to a 1 nmi circular orbit in an affordable, responsive manner. A liquid rocket powered reusable booster is used to propel the upperstage system with payload to a Mach 7 staging condition. After staging, the reusable booster is designed to fly back to the launch site under air-breathing turbofan propulsion. Meanwhile, the two stage upperstage carries the payload to an initial elliptical parking orbit, and eventually inserts the payload in the desired circular orbit. Fig. 2 illustrates the concept of operations (CONOPS) for the SEI ARES Hybrid-OS. 3

Upperstage staging event 21.6 kft/s (relative), 421 kft Booster staging event Mach 7, 167 kft 15 klbs payload delivery MECO at 5x1 nmi. orbit Circularize to 1x1 nmi. 28.5 degree inclination Booster RTLS using turbine engines at Mach.6 and 23 kft Approximately 32 nmi. downrange Liftoff from Cape Canaveral, FL Military Space Port T/W = 1.3, GLOW = 71,75 lbs Figure 2. Concept of Operations (CONOPS) for the SEI ARES Hybrid-OS. For the purposes of this study, it was assumed that the launch site for the SEI ARES Hybrid-OS is a future military spaceport located at Cape Canaveral, Florida. The concept would require a new vertical launch facility at the Cape, in addition to support and processing facilities. Detailed analysis of potential infrastructure and ground operations issues was beyond the scope of this study, however. III. Baseline Concept Results Summary The SEI ARES Hybrid-OS conceptual design study lasted for approximately five months from initial kickoff to final review. A number of design revisions occurred over the course of the project, and consequently it is useful to provide some background information before presenting the results of the disciplinary analysis and trade studies. A. Convergence History It is possible to identify three key design points in the evolution of the SEI ARES Hybrid-OS concept. The first design point represents the initial closed Hybrid-OS that was released on about July 12, 25, 1.5 months into the project. The second point corresponds to the state of the design circa August 22, or about 2.5 months into the five month project. Two significant design changes between the first and second closure points were the temporary elimination of all TPS on the booster, and the addition of a base drag component to the booster flyback aerodynamics model. The third and final point represents the state of the design as presented at the project final review meeting at the Air Force Research Laboratory (AFRL) at Wright-Patterson Air Force Base on March 23, 26. Changes to the design between points two and three included redesigning the upperstage payload fairing to accommodate a representative 15, lb payload, and reassessing the flyback range, thrust required, and engine performance. This third point is considered to be the baseline SEI ARES Hybrid-OS concept resulting from this design study. Figure 3 illustrates the change in gross liftoff weight and upperstage gross weight over the course of the design study. 4

775, 15, 765, 755, System Upperstage 145, 14, System Gross Weight (lbm) 745, 735, 725, 715, 75, 135, 13, 125, 12, 115, Upperstage Gross Weight (lbm) 695, 11, 685, 15, 675, 1 2 3 Closure Point 1, Figure 3. SEI ARES Hybrid-OS gross weight convergence history. B. Final Concept Closure Results The baseline SEI ARES Hybrid-OS booster measures 15 feet in length, and has a wingspan of 55 feet and a tail height of 24 feet with landing gear extended. The two-stage upperstage system weighs 6,82 lb before fueling and without a payload. Including propellants and a 15, lb payload, the upperstage has a gross weight of 141,76 lb. Meanwhile, the SEI ARES Hybrid-OS booster has a dry weight of 78,87 lb, and a gross liftoff weight (GLOW) of 71,75 lb when mated to the upperstage and filled with propellant. A 3-view of the baseline SEI ARES Hybrid-OS booster concept is provided in Fig. 4. Table 1 lists key performance metrics for the booster stage. Figure 5 and Table 2 respectively, show a drawing of the upperstage system and key performance specifications for the second and third stages. 24 feet 15 feet 55 feet Figure 4. SEI ARES Hybrid-OS baseline outer moldline layout and geometry. 5

Table 1. SEI ARES Hybrid-OS booster key performance metrics. Parameter Value Propellants LOX / RP-1 System GLOW (lb) 71,75 Booster Dry Weight (lb) 78,87 Booster Gross Weight (lb) 588,98 Ascent Mass Ratio 2.96 Flyback Mass Ratio 1.15 Ascent Mixture Ratio 2.7 A - A B - B C - C 11 feet r = 3.7 feet r = 4.17 feet 96 feet A B C A B C Figure 5. SEI ARES Hybrid-OS baseline Falcon-1 derived upperstage system. Table 2. SEI ARES Hybrid-OS upperstage key performance metrics. Parameter Value Stage Gross Weight (lb) 141,76 2 nd Stage Dry Weight (lb) 6,2 3 rd Stage Dry Weight (lb) 62 Payload (lb) 15, 2 nd Stage Mass Ratio 4.63 3 rd Stage Mass Ratio 1.32 2 nd Stage Mixture Ratio 2.17 3 rd Stage Mixture Ratio 2.35 The following table and figure summarize the performance of the overall ARES Hybrid-OS launch architecture. Table 3 provides a breakdown of the delta-v losses experienced by the Hybrid-OS as it ascends to orbit, along with a calculated I* value for both the system as a whole and the individual stages. Meanwhile, Fig. 6 illustrates the relative contributions of flight Delta-V and losses to the total Delta-V required for the entire launch system. 6

Table 3. Summary of velocity losses, delta-v, and I* for the SEI ARES Hybrid-OS. Parameter Booster 2 nd Stage 3 rd Stage System Gravity Losses (fps) 3,368 755 263 4,386 Drag Losses (fps) 453 6 459 Atmospheric Thrust Losses (fps) 243 243 TVC Losses (fps) 15 19 3 127 Flight Delta-V (fps) 7,519 14,144 2,527 24,19 Total Delta-V (fps) 11,598 15,14 2,793 29,45 I* (seconds) 215.2 286.6 282.7 259.5 Gravity Losses 14.9% TVC Losses.4% Drag Losses 1.6% Atmospheric Thrust Losses.8% Flight Delta-V 82.3% Figure 6. SEI ARES Hybrid-OS relative contributions of flight delta-v and losses to total delta-v required. IV. Design Process and Methodology A number of industry standard and SEI-developed analysis tools were used to design the baseline ARES Hybrid- OS concept vehicle. Figure 7 is a diagram illustrating the integration of these disciplinary analysis tools in the context of the design process. At the core of this process is an iterative loop between the weights and sizing, ascent trajectory, and flyback trajectory models. Here, the booster and upperstage system are resized to converge the system with mass ratio requirements from the trajectory tools. Several tools including CAD / solid modeling, aerodynamics, and booster propulsion are executed to generate scalable values that initially feed into the iterative convergence loop. Subsequent to the convergence loop life cycle cost, operations, and safety and reliability tools are executed using the details of the closed vehicle concept. The results of the life cycle cost, operations, and safety and reliability tools are not included in this technical paper and will instead be featured in a future publication. 7

CAD / Solid Modeling Feed-forward Links Solid Edge Aerodynamics APAS / SHABP Booster Propulsion REDTOP-2 Weights & Sizing Excel MERs Ascent Trajectory POST II Flyback Trajectory Flyback-Sim Life Cycle Cost NAFCOM Operations Feedback Links AATE / FGOA Safety and Reliability Figure 7. Design structure matrix (DSM) for SEI ARES Hybrid-OS concept. Specific industry codes used in the design of the ARES Hybrid-OS include: GT-SAFETY II POST 1 and 2 APAS - S/HABP Solid Edge NAFCOM AATe 3-degree of freedom trajectory simulation code aerodynamic and aeroheating load predictions configuration layout and center-of-gravity assessments research, development, test, & evaluation (RDT&E) and acquisition costs operations modeling Tools developed by SEI in-house and used in the design of the ARES Hybrid-OS include: REDTOP-2 Excel MERs Sentry SESAW GT Safety-II FGOA liquid rocket engine design (booster rockets and upperstage engine) airframe, tank, and subsystem weight modeling and system sizing transient, 1-D aeroheating analysis and thermal protection system design/sizing for booster stage avionics subsystems vehicle safety and reliability metrics facilities and ground equipment modeling V. Disciplinary Results A. Trajectory Analysis The trajectory of the SEI ARES Hybrid-OS concept has been simulated using POST-2, and is comprised of several distinct phases, each bounded by important mission events. 2 At launch, all elements of the Hybrid-OS, including the reusable booster, upperstage, and payload comprise a single, mated system. The mated system is designed to launch from a vertical orientation, supported by an unspecified pad structure. When the three main liquid rocket engines on the booster stage ignite for liftoff, the system has a thrust-to-weight ratio of 1.3. Within 1 seconds of liftoff, the ARES Hybrid-OS will be more than 5 feet above the launch pad traveling at a relative velocity of over 1 fps. At T + 6 seconds, the vehicle will have surpassed an altitude 23, feet and a relative velocity of over 1, fps. Just over two and a half minutes into the flight the first staging point will occur at an altitude of 167, feet and a Mach number of 6.9. Here the upperstage system and payload will separate from the booster and the 2nd stage main rocket engines will be ignited. The SEI ARES Hybrid-OS 2nd stage burns for approximately 19 seconds as it accelerates to orbit. By the time the 2nd stage shuts down, the vehicle is traveling at a relative velocity of over 21, fps at an altitude of over 8

421, feet. After the 2nd stage is jettisoned, the 3rd stage and payload coast for a period of about 1 seconds in a suborbital arc before the 3rd stage engine is fired to accelerate the payload into the desired circular low Earth orbit. Meanwhile, as the upperstage system is ascending to orbit, the ARES Hybrid-OS booster begins maneuvering to return to the launch site. A short time after separation of the upperstage, the booster enters an unpowered turn and descent maneuver to attain the desired flight path heading. Two low-bypass turbofan engines are engaged once the booster stage has decelerated to subsonic velocity and descended to just under 3, feet altitude. At Mach.6, the flyback cruise lasts about 1 hour after the booster and upperstage separate. In other words, the ARES Hybrid-OS booster would arrive back at the launch site just a few minutes over an hour after launch. The following figures illustrate some of the most significant results of the trajectory analysis for the SEI ARES Hybrid-OS. Figure 8 plots the Mach number of the booster, as well as the altitude of both the booster and upperstage versus the flight time. After the trajectory branches, the upper curve represents the upperstage, while the lower curve represents the booster stage. Note that this plot has been truncated at 1, seconds and, thus, excludes most of the lengthy booster flyback segment of the mission. Figure 9 shows the relative flight velocity of the ARES Hybrid-OS versus mission time from launch until orbit insertion. The data in Fig. 9 corresponds to the ascent trajectory only and can be thought of as displaying the velocity of the payload as each stage contributes delta-v on the way to orbit. Finally, the dynamic pressure experienced by the ARES Hybrid-OS is plotted along with altitude versus the mission time in Fig. 1. The maximum dynamic pressure experienced by the booster is about 72 psf at T+7 seconds. 8 64, 3, Booster Mach Number 7 6 5 4 3 2 1 Mach Number Altitude 48, 32, 16, Altitude (ft) Relative Flight Velocity (fps). 25, 2, 15, 1, 5, 2 4 6 8 1, Time (s) Figure 8. SEI ARES Hybrid-OS flight Mach number and altitude vs. time. 8 1 2 3 4 5 6 7 Time (s) Figure 9. SEI ARES Hybrid-OS ascent flight velocity vs. time. 64, 7 Dynamic Pressure (psf) 6 5 4 3 2 Dynamic Pressure Altitude 48, 32, 16, Altitude (ft) 1 1 2 3 4 5 6 7 Time (s) Figure 1. SEI ARES Hybrid-OS ascent freestream dynamic pressure and altitude vs. time. B. Propulsion The SEI ARES Hybrid-OS concept employs several propulsion systems not only to boost a payload into Earth orbit, but also to return the reusable booster stage to the launch site. The Hybrid-OS booster incorporates an all-new engine design, while the upperstage relies on two different in-production engine models. Based on study groundrules, all liquid rocket engines on the Hybrid-OS burn liquid oxygen and hydrocarbon rocket fuel (RP-1). 9

Figure 11 contains a plot of thrust versus time for the ascent portion of the Hybrid-OS trajectory. From this figure one can see the transition of the booster engine from sea level static (SLS) thrust to near-vacuum performance, as well as the vacuum thrust levels of the 2nd and 3rd stage rocket engines. Figure 11 does not address the thrust level of the turbofan engines during the flyback cruise. Figure 12 shows similar performance trends for the specific impulse of each of the ascent rocket engine systems. Thrust (lb f ) 1,2, 1,, 8, 6, 4, 2, 1 2 3 4 5 6 7 Time (s) Figure 11. SEI ARES Hybrid-OS thrust vs. time. Net Isp (s) 4 35 3 25 2 15 1 5 1 2 3 4 5 6 7 Time (s) Figure 12. SEI ARES Hybrid-OS net specific impulse vs. time. 1. Booster Ascent Engines The SEI ARES Hybrid-OS is powered on ascent by liquid rocket engines located on the reusable booster. Given the assumed design requirement of a vehicle thrust-to-weight ratio of 1.3 at liftoff, as well as the requirement to use a liquid oxygen and RP-1 propellant combination, it is clear that very few existing rocket engines would be suitable for this application. Among existing engines, two designs that could be considered for this application are the NK-33 and RD-18, both of which have a heritage in the former Soviet Union. Table 4 indicates the relevant performance specifications for both of these engines. From this table it can be seen that multiple NK-33 engines would be required to achieve a liftoff thrust-to-weight of at least 1.3 for a launch vehicle gross weight of just over 7, lb, while a single RD- 18 unit is nearly adequate for the task. Table 4. Existing LOX/RP engine performance specifications. Parameter NK-33 RD-18 Cycle Staged-Combustion Staged-Combustion Turbine Flow Single Shaft Single Shaft Propellants LOX / RP-1 LOX / RP-1 Mixture Ratio 2.6 2.6 Chamber Pressure (psia) 2,19 3,734 Area Ratio 27:1 36.4:1 Thrust-to-Weight Ratio @ SLS / Vacuum 124 / 152 8 / 87 Thrust @ SLS / Vacuum (lbf) 339, / 413,61 86,4 / 933, Isp @ SLS / Vacuum (seconds) 297 / 331 311 / 337 Exit Area (ft 2 ) 18.79 35.32 Minimum Length (feet) 12.2 13. Although one or both of these existing LOX / RP-1 engine designs may have been well-suited to the SEI ARES Hybrid-OS system, a ground rule was established which stated that only domestic technologies would be used on the vehicle concept. Therefore, in the absence of any applicable off-the-shelf domestic engine designs it was assumed that the development of an all-new LOX / RP-1 engine would be undertaken to enable the ARES Hybrid-OS. The design requirements for this all-new engine design are listed in Table 5. A conceptual illustration of the engine is provided in Fig. 13. 1

Table 5. SEI ARES Hybrid-OS booster all-new LOX/RP-1 engine design requirements. Parameter All-new LOX / RP-1 Engine Design Value Thrust @ SLS Re-sized as necessary to provide vehicle thrust-to-weight of 1.3 at launch (i.e. rubberized engine design) Isp @ SLS (seconds) 35 Thrust-to-Weight Ratio @ SLS 15:1 Figure 13. Conceptual illustration of all-new LOX/RP booster engine. The SEI ARES Hybrid-OS concept was designed with three all-new engines on the booster stage sized to provide the required lift-off thrust-to-weight when operated at 1% throttle. There was no capability for engine out operation designed into the concept. 2. Booster Attitude Control System (ACS) The booster stage of the SEI ARES Hybrid-OS concept is equipped with an attitude control system (ACS) to augment post-cutoff aerodynamic control surfaces as the booster passes through a high altitude, low dynamic pressure arc after staging. The ACS consists of forward and aft thruster groups on both sides of the booster. In keeping with the desire for efficient ground operations and responsiveness, a propellant combination of gaseous oxygen (GOX) and ethanol was baselined. This system was assumed to possess an Isp of 28 seconds and was sized to provide a total Delta-V of 5 fps. 3. Booster Flyback Engines In order to reuse the ARES Hybrid-OS booster it must return to a military base either on its own power, or with the assistance of another vehicle. For this study it was assumed that the booster would use low-bypass turbofan engines to return to the site from which it was launched. A flyback cruise altitude of 23, feet at a Mach number of.6 was selected based on a series of trade studies. This flyback trajectory requires a total sea level static (SLS) thrust of 38,19 lbf to be provided by the jet engines. The General Electric F118 engine is rated at 19, lbf thrust SLS without augmentation, and therefore would be sufficient to meet this performance requirement in a parallel installation. 3 Additional specifications for the F118-GE- 1 engine are provided in Table 6. Although this study focused on low-bypass turbofan flyback engines, other configurations or flyback approaches could also be explored in future studies. 11

Table 6. F118-GE-1 low-bypass turbofan engine specifications. Parameter Value Fan / Compressor Stages 3 / 9 Low-Pressure Turbine Stages/ High-Pressure Turbine Stages 2 / 1 Maximum Diameter (inches) 46.5 Length (inches) 1.5 Dry Weight (lb) 3,2 Specific Fuel Consumption @ Max Power (lb/lb-hr).67 Max SLS Thrust Level (lbf) 19, Overall Pressure Ratio at Max Power 35:1 4. Upperstage Engines The ARES Hybrid-OS concept examined in this study incorporated a two stage upperstage, both using a combination of liquid oxygen (LOX) and rocket propellant (RP-1) as oxidizer and fuel, respectively. The rocket propulsion technology used on each element of the upperstage consists of existing Space Exploration Technologies (SpaceX) engine designs. Namely, the SEI ARES Hybrid-OS second stage is powered by two SpaceX Merlin engines, while the third stage is powered by a single SpaceX Kestrel engine. 4 The Merlin engine is a new LOX / RP-1 engine developed by SpaceX to power the first stage of the Falcon-1 launch vehicle. The engine uses a gas-generator cycle and a pintle style injector, and produces about 92, lbf of thrust in a vacuum. Also a new design, the Kestrel is a pressure-fed engine with a pintle style injector. The Kestrel is sized to produce 7, lbf of thrust in a vacuum. Detailed specifications for both the Merlin and Kestrel engines can be found in Table 7 below. Figure 14 contains CAD renderings of the Merlin and Kestrel engines. Table 7. SpaceX Merlin and Kestrel engine specifications. 5 Parameter Merlin Kestrel Cycle Gas Generator Pressure-Fed Configuration Single-shaft Turbopump -- Propellants LOX / RP-1 LOX / RP-1 Thrust @ SLS / Vacuum (lbf) 77, / 92, -- / 7, Mixture Ratio 2.17 2.35 Chamber Pressure (psia) 766 15 Area Ratio 14.5:1 6:1 Thrust-to-Weight Ratio @ SLS 96 42 Isp @ SLS / Vacuum (seconds) 255 / 34 -- / 327 Figure 14. SpaceX Merlin (left) and SpaceX Kestrel (right) used on the SEI ARES Hybrid-OS 2 nd and 3 rd stages, respectively. Note: images not to scale. 12

C. Weights and Sizing (W&S) Weights and sizing analysis for the SEI ARES Hybrid-OS concept was accomplished using physical and historical mass estimating relationships (MERs) encapsulated in a Microsoft Excel workbook. Each stage of the Hybrid-OS was represented by a parametric sizing model designed to accept a required mass ratio from the trajectory analysis and to return a corresponding stage geometry and estimated weight breakdown. An iterative loop was completed between the weights and sizing models and the trajectory analysis in order to yield a closed vehicle concept. Of particular interest during the creation of the weights and sizing models for the Hybrid-OS was the desire to validate the upperstage weight estimating approach with actual Space Exploration Technologies (SpaceX) Falcon-1 data. A bottom-up analysis of subsystem weight estimates was used to generate in-house estimates of the vehicle dry and gross weights for comparison with the actual Falcon-1. Recall that the 2nd stage of the Hybrid-OS concept is derived from the 1st stage of the Falcon-1, while the 3rd stage of the Hybrid-OS is based on the Falcon-1 2nd stage. Figure 15 gives a comparison of the existing SpaceX Falcon-1 launch vehicle with the new Hybrid-OS upperstage. Material selection, subsystem technologies, and propulsion systems for the Hybrid-OS upperstage are assumed to be in line with SpaceX state-of-the-art in these areas. 96 feet 7 feet Parameter Dry Weight (lb) Gross Weight (lb) Length (feet) Fairing Diameter (feet) SpaceX Falcon-1 3,66 6, 7 4.9 Hybrid-OS Upperstage 6,82 141,76 96 8.3 Figure 15. Comparison of SpaceX Falcon-1 vehicle with SEI ARES Hybrid-OS upperstage. 1. Booster Sizing and Mass Properties Table 8 indicates the estimated weight breakdown for the baseline SEI ARES Hybrid-OS booster stage. The most significant portion of the stage dry weight comes from structural elements including both aerodynamic surfaces and the primary airframe. Installed propulsion systems make up the second largest dry weight line item. As indicated, a 15% dry weight margin or manager s reserve was applied to this stage. 13

Table 8. SEI ARES Hybrid-OS booster level-1 hardware and fluid weights. Hardware Element Weight (lb) Fluids Weight (lb) Wings and Tails 12,61 Residuals 2,39 Airframe (aeroshell, tanks, bulkheads, etc.) 19,23 Reserves 4, Thermal Protection Systems 39 Flyback/RTLS 12,4 Landing Gear 2,44 Attitude Control System (ACS) 56 MPS 24,2 Unusables 47 Liquid Rocket Engines 15,86 Ascent Fuel (RP-1) 127,11 Turbofan Engines 8,16 Ascent Oxidizer (LOX) 343,18 Attitude Control System (ACS) 1,31 Total All Propulsive Fluids 49,11 Subsystems (avionics, EPC&D, EHAs, etc.) 8,58 Dry Weight Margin (15%) 1,29 Total Hardware 78,87 Total Hardware + Fluids + Upperstage 71,75 2. Upperstage Sizing and Mass Properties Table 9 and Table 1 contain the weight breakdown estimates for the SEI ARES Hybrid-OS 2nd stage and 3rd stage, respectively. A smaller 5% dry weight margin was applied to these estimates in recognition of the extensive use of existing systems and technologies. The SpaceX Falcon-1 launch vehicle is designed for a nominal payload of 1,25 lb to low Earth orbit. Given the substantial difference between this value and the ARES objective payload of 15, lb, it was necessary to design an all-new aerodynamic fairing for the upperstage of the Hybrid-OS launch vehicle. Table 11 contains the sizing assumptions and results of this design. Table 9. SEI ARES Hybrid-OS 2 nd stage level-1 hardware and fluid weights. Hardware Element Weight (lb) Fluids Weight (lb) Wings and Tails Residuals 56 Airframe (aeroshell, tanks, bulkheads, etc.) 3,17 Reserves 1,11 Thermal Protection Systems Attitude Control System (ACS) Landing Gear Unusables MPS 2,71 Pressurant Gas 2 Merlin Liquid Rocket Engines (x2) 2,71 Ascent Fuel (RP-1) 35,6 Attitude Control System (ACS) Ascent Oxidizer (LOX) 76,8 Subsystems (avionics, EPC&D, EHAs, etc.) 2 Total All Fluids 112,83 Dry Weight Margin (5%) 3 Total Hardware 6,2 Total Hardware + Fluids + Payload Fairing 121,3 14

Table 1. SEI ARES Hybrid-OS 3 rd stage level-1 hardware and fluid weights. Hardware Element Weight (lb) Fluids Weight (lb) Wings and Tails Residuals Airframe (aeroshell, tanks, bulkheads, etc.) 31 Reserves 5 Thermal Protection Systems Attitude Control System (ACS) Landing Gear Unusables MPS 21 Pressurant Gas Kestrel Liquid Rocket Engine (x1) 21 Ascent Fuel (RP-1) 1,51 Attitude Control System (ACS) Ascent Oxidizer (LOX) 3,55 Subsystems (avionics, EPC&D, EHAs, etc.) 7 Total All Fluids 5,11 Dry Weight Margin (5%) 3 Total Hardware 62 Total Hardware + Fluids + Payload 2,73 Table 11. SEI ARES Hybrid-OS upperstage payload fairing sizing data. Parameter Value Assumed Payload Density (pcf) 1. Fairing Diameter (feet) 8.33 Fairing Length (feet) 3 Material Type Aluminum Total Fairing Mass (lb) 2, Total Fairing Volume (ft 3 ) 1,5 D. Aerodynamics The SEI ARES Hybrid-OS booster is a wing-body configuration featuring a double delta wing, twin canted vertical stabilizers, and a canard mounted above the flyback engines. The size and shape of the Hybrid-OS fuselage are driven by the underlying propellant tank configuration. Landing and subsonic cruise play a role in the design of the lifting surfaces, although the demands of traveling at supersonic and hypersonic speeds are the predominant influence. The theoretical wing planform area of the Hybrid-OS booster is 1,395 ft 2, with a wingspan of 55 feet. The double delta wing features an inner sweep angle of 62 and an outer sweep angle of 45. A NACA-65 series airfoil has been selected for use not only on the main delta wing, but also on the tails and the canards. Table 12 lists some pertinent aerodynamic design data for the ARES Hybrid-OS. Table 12. SEI ARES Hybrid-OS aerodynamic surface specifications. Parameter Wings Tails Canards NACA Airfoil Type 65-Series 65-Series 65-Series Span / Height (feet) 55 12 12 Sweep Angle ( ) 62 / 45 6 3 Exposed Planform Area (ft 2 ) 1,1 265 11 Aerodynamic coefficients were calculated for the ARES Hybrid-OS configuration using the Aerodynamic Preliminary Analysis System (APAS). 6 Figure 16 is a top view of the surface node-based geometry that was modeled in APAS. Figure 17 shows lift and drag coefficient data generated by APAS for a particular angle of attack over a range of Mach numbers. Figure 18 is a plot of as-flown angle of attack from the ARES Hybrid-OS ascent trajectory simulation. 15

Figure 16. Aerodynamic analysis grid for SEI ARES Hybrid-OS..35 1 8 C L, C D.3.25.2.15.1 CL CD Angle of Attack (degrees) 8 6 4 2-2 Angle of Attack Dynamic Pressure 7 6 5 4 3 2 Dynamic Pressure (psf).5-4 1. 2 4 6 8 1 Mach Number Figure 17. Lift and drag coefficients for SEI ARES Hybrid-OS booster at AOA=5. -6 1 2 3 4 5 6 7 Time (s) Figure 18. SEI ARES Hybrid-OS as-flown angle-ofattack vs. time. E. Aeroheating and Thermal Protection A fundamental tenet of the ARES design philosophy is the idea that eliminating the need for an extensive thermal protection system on the reusable booster stage will yield great benefits to affordability, operability, and responsiveness. To this end, it has been proposed that the staging point between the booster and upperstage be pushed back to a lower Mach number than might normally be selected in a traditional approach. SEI analyzed the heat rate over the course of the booster ascent and flyback trajectories, as well as the distribution of maximum surface temperatures on the Hybrid-OS concept vehicle. The results of these two analyses are shown in Fig. 19 and Fig. 2, respectively. The conclusion to be drawn from these data is that the vast majority of the booster skin structure will not need additional thermal protection. Some particularly hot regions such as the nose and wing/tail leading edges will require limited TPS, however. 16

Booster Heat Rate (Btu. / ft 2 - s) 12 1 8 6 4 2 Booster Heat Rate Altitude 2 4 6 8 1, Time (s) 3, 25, 2, 15, 1, 5, Altitude (ft) Windward Leeward Max T surface (R) 84 7 56 Figure 19. SEI ARES Hybrid-OS booster heat rate and altitude vs. time. Figure 2. SEI ARES Hybrid-OS booster maximum surface temperature distribution. SEI s Sentry analysis code was used to determine approximate surface temperatures for the Hybrid-OS booster. 7 Convective heat rates were provided for use in Sentry by S/HABP. It should be noted that the analysis performed at each node was one dimensional, and therefore significant airframe conduction was not accounted for. Essentially this implies that the high temperatures on the windward side of the booster are likely overpredicted, while the cool temperatures on the leeward side are likely underpredicted. VI. Observations and Conclusions Based on the results of this conceptual design activity, SEI has been able to answer several key questions regarding the feasibility and viability of this ARES Hybrid-OS concept. For instance, the trajectory simulations and aeroheating analysis performed as part of the design closure reveal that it would be possible to forego an extensive thermal protection system on the booster one of the linchpins of the highly operable ARES approach. SEI s analysis also explored the potential synergy between low cost commercial small payload launch vehicles and the design requirements for the hybrid system s expendable upperstage. In addition, a detailed examination of the ARES Hybrid-OS booster main propulsion elements enabled SEI to refine previously vague design guidelines for development of these liquid locket engines, including various material trade studies. In the context of the analysis presented here, it is evident to the authors that the trajectory of the SEI ARES Hybrid-OS is suboptimal in terms of traditional performance metrics (gross mass, etc.). In other words, from the point of view of system gross mass, a more advantageous design point would call for a higher Mach number staging point between the booster and the upperstage. However, the conclusion offered here is that the potential operational benefits of maintaining a lower Mach number staging condition such that an extensive TPS is not required may be worth the performance sacrifice. It was discovered during the course of this design study that the characteristics of the booster flyback system are significant drivers in the performance of the overall system. While SEI examined using jet engines to power a flyback trajectory, future studies should consider rocket boost-back and other alternatives. Acknowledgments The authors would like to recognize the contribution of John Olds of SpaceWorks Engineering, Inc. to the design of the SEI ARES Hybrid-OS concept vehicle. The authors would like to thank Eric Paulson of AFRL/PRST at Edwards AFB for his comments and suggestions regarding the design of the booster liquid rocket engines. The authors would like to thank Chuck Bauer of Universal Technology Corporation for his contributions to the turbine engine system selection. The authors would also like to express their sincere appreciation to Glenn Liston from the Air Force Research Laboratory at Wright-Patterson Air Force Base for funding this work under Contract #F33615-3-C-232. 17

References 1. FALCON: Force Application and Launch from CONUS, Task 1 Small Launch Vehicle (SLV) Phase II, Program Solicitation Number 4-5, Defense Advanced Research Projects Agency, Arlington, VA, 24. 2. POST II, Program to Optimize Simulated Trajectories, Software Package, Ver. 1.1.6.G, NASA Langley Research Center, Hampton, VA, 24. 3. General Electric, GE - Aviation: F118, [online], URL: http://www.geae.com/engines/military/f118/index.html [cited April 26] 4. Space Exploration Technologies Corporation, Falcon Engines, [online], URL: http://www.spacex.com [cited: October 26] 5. Isakowitz, S.J., J.B. Hopkins, J.P. Hopkins, International Reference Guide to Space Launch Systems, Fourth Edition,, Reston, Virginia, 24. 6. Sova, G., Divan, P., Aerodynamic Preliminary Analysis System II, Part II User s Manual, NASA CR 18277, April, 1991. 7. Bradford, J.E., Olds, J.R., Thermal Protection System Sizing and Selection for RLVs Using the Sentry Code, AIAA 26-465, 42 nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Sacramento, CA, 9-12 July 26. 18