IAC-02-S.5.02 XCALIBUR: A VERTICAL TAKEOFF TSTO RLV CONCEPT WITH A HEDM UPPERSTAGE AND A SCRAM-ROCKET BOOSTER J. E. Bradford A. Charania J. R. Olds M. Graham SpaceWorks Engineering, Inc. (SEI) Atlanta, GA U.S.A.

53rd International Astronautical Congress

The World Space Congress - 2002 10-19 Oct 2002/Houston, Texas

For permission to copy or to republish, contact the International Astronautical Federation 3-5 Rue Mario-Nikis, 75015 Paris, France

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XCALIBUR: A VERTICAL TAKEOFF TSTO RLV CONCEPT WITH A HEDM UPPERSTAGE AND A SCRAM-ROCKET BOOSTER

John E. Bradford†, A.C. Charania*, John R. Olds±, and Matthew Graham§

SpaceWorks Engineering, Inc. (SEI) Atlanta, GA U.S.A.

www.sei.aero

ABSTRACT A new 3rd generation, two-stage-to-orbit (TSTO) reusable launch vehicle (RLV) has been designed. The Xcalibur concept represents a novel approach due to its integration method for the upperstage element of the system. The vertical-takeoff booster, which is powered by rocket-based combined-cycle (RBCC) engines, carries the upperstage internally in the aft section of the airframe to a Mach 15.5 staging condition. The upperstage is released from the booster and carries the 20 Klbs of payload to low earth orbit (LEO) using its high energy density matter (HEDM) propulsion system. The booster element is capable of returning to the original launch site in a ramjet-cruise propulsion mode. Both the booster and the upperstage utilize advanced technologies including: graphite-epoxy tanks, metal-matrix composites, UHTC TPS materials, electro-mechanical actuators (EMAs), and lightweight subsystems (avionics, power distribution, etc.). Details of the concept design including external and internal configuration, mass properties, engine performance, trajectory analysis, aeroheating results, and economics assessment are given. Highlights of the distributed, collaborative design approach, analysis tools, and a summary of a payload trade study are also provided.

NOMENCLATURE AAR air-augmented rocket AATE Architecture Assessment Tool-Enhanced

CABAM Cost and Business Analysis Module CER cost-estimating relationship COTS commercial off-the-shelf CSTS Commercial Space Transportation Study Ct thrust coefficient DDT&E design, development, test, & evaluation EMA electro-mechanical actuators ETO earth to orbit FY fiscal year GLOW gross lift-off weight HEDM high energy density matter IOC initial operating capability IRR internal rate of return Isp specific impulse (sec.) I* equivalent trajectory averaged Isp (sec.) KSC Kennedy Space Center LCC life cycle cost LEO low earth orbit (100-150 nmi. altitude) LH2 liquid hydrogen LOX liquid oxygen LRU line replacement unit MER mass estimating relationship MMC metal matrix composites MR mass ratio (gross weight/burnout weight) NAFCOM NASA/Air-Force cost model NPV net present value RBCC rocket-based combined-cycle RLV reusable launch vehicle SCORES SpaceCraft Object-oriented Rocket

Engine Simulation SCCREAM Simulated Combined Cycle Rocket

Engine Analysis Module SEI SpaceWorks Engineering, Inc. SLS sea-level static SR scram-rocket TCAT Thermal Calculations Analysis Tool TFU theoretical first unit TPS thermal protection system Q dynamic pressure (psf) T/We installed engine thrust-to-weight UHTC ultra high temperature ceramic

_____________________________________________ † - Director of Hypersonics, Member AIAA * - Senior Futurist, Member AIAA

±- President, Associate Member AIAA § - Director of Concept Development Copyright ©2002 by SpaceWorks Engineering, Inc. (SEI). Published by the American Institute of Aeronautics and Astronautics, Inc. Released to IAF/IAA/AIAA to publish in all forms.

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INTRODUCTION NASA Marshall Space Flight Center’s Advanced Space Transportation Program (ASTP) is currently conducting a number of system analysis studies focused on a variety of combined-cycle propulsion system options for use on future (2020+) reusable launch vehicles1,2. One specific option currently under consideration is the rocket-based combined cycle (RBCC) engine. These multi-mode engines combine the best aspects of rocket propulsion (high thrust-to-weight) and airbreathing propulsion (high specific impulse). Previous research has shown that vehicles utilizing RBCC propulsion are attractive candidates for future space access due to the variety of missions and markets they can support (ETO payload delivery, military global reach, space tourism, etc.)3,4. While a number of conceptual design organizations have focused their research on single-stage and two-stage, horizontal takeoff concepts, there appear to be very few studies examining two-stage vertical takeoff with horizontal landing systems utilizing RBCC engines. This is likely attributable to the difficulties of arranging the two vehicles vertically into the airframe/propulsion integrated configuration required for hypersonic air-breathing propulsion operation. With this clearly identifiable unexplored region of the design space, SpaceWorks Engineering, Inc. (SEI) was tasked by the Advanced Concepts Division (TD30) at NASA MSFC to undertake the conceptual-level vehicle study of a two-stage vertical 3rd Gen RLV using RBCC propulsion on the booster stage. The Xcalibur concept was thus created and designed by SEI and represents a potentially feasible 3rd Gen RLV concept in this undeveloped region of the design space. The Xcalibur concept consists of two elements: the booster and the upperstage. The booster component is enabled by its main propulsion system that utilizes four RBCC engines. These engines operate in three distinct modes: air-augmented rocket (AAR), ramjet, and scram-rocket (SR). The engines operate in AAR mode from takeoff to Mach 2.5, with ramjet mode operation from Mach 2.5 to Mach 5.5. The full-throttle thruster re-ignition for SR mode occurs at Mach 5.5 and extends to the staging condition at approximately Mach 15.5. The extended utilization of the scram-

rocket mode greatly improves vehicle performance by providing superior acceleration when compared to the scramjet mode performance over the same flight regime. Results indicate that the specific impulse penalty due to the scram-rocket mode operation is outweighed by the greater acceleration, reduced flight time, smaller vehicle size due to increased mixture ratio, and lower allowable dynamic pressures (q). A complete vehicle life-cycle analysis was performed in a multi-disciplinary design environment for this concept. Disciplinary performance analysis tools that were automated include: ascent trajectory (POST), RBCC propulsion (SCCREAM), liquid rocket propulsion (SCORES-II), booster flyback simulation with Excel spreadsheet, and an Excel model for component weight estimation. These tools were automated in Phoenix Integration’s ModelCenter© collaborative design environment. Performance tools utilized for the analysis, but not requiring automation, included the IDEAS-SDRC software for solid modeling, SEI’s TCAT-II for aeroheating/TPS sizing, and APAS-S/HABP for the aerodynamic analysis. The paper describes the vehicle concept and operation, discussing the types of technologies used and the nominal flight scenario. A brief discussion explaining the decision-making process for the vehicle configuration is included. For cost predictions, NASA Air Force Cost Model (NAFCOM) derived cost estimating relationships (CERs) were used. Other metric predictions were developed using a number of codes, including SEI’s CABAM_A (financials), NASA KSC’s AATe (operations), and Georgia Institute of Technology’s GT-SafetyII (safety and reliability). All of the above metric assessment tools were also wrapped and automated in the ModelCenter© environment.

THE XCALIBUR CONCEPT Overview Xcalibur is a two-stage, fully reusable vehicle consisting of booster and upperstage elements (see Figure 1). The booster geometry consists of an initial conical shape with a short transition to 2-D forebody at the main engine inlet face. The four LOX/LH2 RBCC engines are sidemounted on the booster, with two engines per side.

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The vertical takeoff arrangement allows for smaller wings and reduced undercarriage weight. The side mounted engines permit the thrust vector to be directed through the vehicle center of gravity and minimizes the propulsion pitch-plane moment common in other underslung engine arrangements.

Figure 1. Xcalibur Concept During Takeoff The upperstage vehicle is carried internally in the aft section of the booster vehicle. The upperstage propulsion system consists of three LOX/hydrocarbon High-Energy Density Matter (HEDM) liquid rocket engines. The HEDM fuel being used is representative of strained-ring hydrocarbons like AFRL-1 or Quadricyclane. Both of these fuels are currently being investigated by NASA and the Air Force. The three upperstage engines are mounted to the vehicle’s base region and are exposed from the booster. In this manner, the upperstage engines can be used during launch to offset the thrust required by the booster to ascend vertically. The upperstage fuselage is a lifting body design as shown in Figures 2 and 3, with fully articulating control surfaces for stability during reenty. Currently, the vehicle design calls for the upperstage tail fins to fold approximately 90o for stowage within the booster.

The baseline vehicle concept is designed to deliver 20,000 lbs. of payload into a 100 nmi. x 28.5° circular orbit from a future spaceport based at the current Kennedy Space Center (KSC) location. The upperstage payload bay can also serve to carry 8 passengers in a human-rated module to the same orbit. The upperstage subsystems are designed to allow the vehicle to remain in orbit autonomously for one day, not including any passenger life support equipment.

Figure 2. Xcalibur Upperstage on Orbit

Figure 3. Upperstage Vehicle after Staging In addition to the RBCC and HEDM propulsion systems, Xcalibur uses a number of advanced technologies currently in various stages of development in NASA’s 2nd Generation and 3rd Generation RLV programs. Lightweight metal matrix composites (MMCs) such as titanium-aluminides are

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used for the primary structure and wings. Graphite composites are used to construct the main propellant tanks in both the booster and upperstage. To avoid active cooling, the high temperature nosecap, wing leading edges and cowl leading edges are constructed of Ultra-High Temperature Ceramic (UHTC) Thermal Protection System (TPS). TUFI TPS tiles and high temperature AFRSI blankets are used to protect the acreage areas of the vehicles. In addition, lightweight power, avionics, and electromechanical surface control actuators are used. Hydraulics are not used on either system. The vehicle is capable of autonomous operation require no pilots. Initial Operational Capability (IOC) is expected to be in the year 2025, with a technology freeze date of 2020. Each vehicle airframe is designed for long life operation, estimated at 1,000 flights per airframe and 500 flights per engine. Mission Profile Xcalibur operates from a notional future spaceport at KSC in Florida, USA. Initial take-off thrust is provided by the AAR mode of the RBCC engines and upperstage HEDM engines. The vehicle is designed for a thrust-to-weight ratio of 1.15 at takeoff with 15% of the total required thrust being provided by the upperstage engines. All the upperstage engines continue firing through transonic to approximately Mach 2, with all the consumed propellant for the upperstage provided by the booster through crossfeed lines. From Mach 2 to 2.5, the vehicle accelerates using the booster RBCC engine AAR mode only and begins following an 1,800 psf constant dynamic pressure (q) boundary profile. At approximately Mach 2.5 the RBCC operating mode is shifted to ramjet operation. There is an initial significant decrease in thrust as the RBCC thrusters are shut down and the ramjet combustor experiences a thermal choke, which requires operating with a fuel flowrate below stoichiometric conditions. This choking is fully relieved by Mach 3.2. The vehicle shifts to scram-rocket mode at Mach 5.5, by decreasing the backpressure to obtain supersonic combustion and restarting the thrusters to a 100% throttle setting. The booster’s scram-rocket engine mode is used to rapidly accelerate the vehicle along the constant-q boundary to Mach 10. At Mach 10, the vehicle pulls off the q-boundary, pitching up and accelerating to a

significantly lower q of 95 psf at the approximately Mach 15.5 staging condition. Upon reaching the staging point, a drag-chute is deployed from the upperstage, pulling the vehicle out the aft section of the booster. Upon separation, the booster performs a bank and dive maneuver, decreasing its altitude and Mach number to a Mach 3 at 80,000 ft ramjet-powered flyback condition. Using the thrust provided by its HEDM propulsion system, the upperstage continues on to an elliptical insertion orbit of at least 30x100 nmi., with MECO at perigee. At the apogee point in the orbit, a small delta-V burn using a single-HEDM engine circularizes the upperstage at the desired 100 nmi. LEO orbit. The payload is then released and the vehicle can be de-orbited for return to KSC.

ENGINEERING DISCIPLINE TOOL OVERVIEW

Modeling helps to determine the properties of a technically feasible design. In the conceptual design stage, modeling can include the use of monolithic synthesis/sizing codes or integrated disciplines in a multi-disciplinary environment. These models are representations of the real world based on processes in terms of physics, human operations, financials, etc. This vehicle study utilizes many engineering level toolsets to design the Xcalibur concept. These disciplines include: trajectory, propulsion, mass properties, aerodynamics, aeroheating, operations, cost, safety/reliability, and economics. Table 1 provides a summary of each discipline and tool used to support the analysis for the respective discipline. Details of each of the engineering disciplines and specific tools utilized for this examination are provided in the following sections.

Table 1. Analysis Tool Summary Discipline Tool/Model/Simulation

CAD SDRC-IDEAS, Solid Edge Aerodynamics APAS Aeroheating TCAT-II Mass Properties MS Excel Worksheets RBCC Propulsion SCCREAM Liquid Rocket Propulsion SCORES-II Ascent Trajectory POST (3 DOF) Flyback MS Excel Worksheet Cost NAFCOM-99 Operations AATe Economics CABAM_A Safety GT-Safety II

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Discipline: Mass Properties The system weights will be determined using a combination of results from higher-level analysis and the standard industry practice of mass estimating relationships (MERs) for subsystems and appropriate average unit weights for vehicle structures. Most of these MERs have a NASA Langley heritage, but were adjusted to account for advanced materials technologies, construction techniques, and lightweight subsystems. For example, MERs were included that estimate the wing weight based on surface area and wing loading. The propellant tank MER was based on design pressure, materials, and internal volume values from prior design studies. Some of the MER values were obtained from other disciplinary analysis conducted for the specific Xcalibur design. The TPS system weights are one example of this type of MER. Vehicle sizing was accomplished through photographically scaling the reference vehicle in the MER-based MS Excel spreadsheet. Specific weight line items are organized in a 28 category, 3-level Weight Breakdown Structure (WBS) for both the upperstage and booster vehicles. For a specified booster and upperstage mass ratios and mixture ratios, the system can be sized by varying the length of each component. Typical two-stage system sizing allows for sizing the upperstage first, then sizing of the booster. The booster sizing does not affect the upperstage. But since the design required maintaining a fixed liftoff Thrust-to-Weight (T/W) ratio and percentage thrust contribution from the upperstage, the two vehicles are coupled. Thus the upperstage and the booster must be sized simultaneously or iteratively to obtain the desired mass ratio and mixture ratio values. This problem can be easily solved from within the MS Excel environment through careful problem formulation. Discipline: Trajectory The trajectory analysis was performed using the three degree-of-freedom version of the Program to Optimize Simulated Trajectories (POST)5. POST is a Lockheed Martin and NASA code that is widely used for trajectory optimization problems in advanced vehicle design. It is a generalized event-oriented code that numerically integrates the equations of motion of a

flight vehicle given definitions of aerodynamic coefficients, propulsion system characteristics, weight models, etc. Numerical optimization is used to satisfy trajectory constraints and minimize a user-defined objective function. The trajectory simulation consisted of an untrimmed analysis of the vehicles starting at time zero to the desired final orbit. Input parameters to the model will include the aerodynamics database with Mach number and angle-of-attack dependency, propulsion system with altitude, Mach number, and engine mode dependency, stage weights, engine mode transition points, and dynamic pressure boundary (maximum q) limits. The simulation is designed to allow an optimizer to control the flight path via the pitch angles and Main Engine Cut-Off (MECO) velocity, with the objective function being to minimize the total consumed weight of both stages. Trajectory constraints include: final perigee and apogee altitudes, minimal change in the pitch rates while entering and exiting the q-boundary path, a maximum wing normal force during SR-mode pull-up, a positive MECO flight path angle, and a maximum acceleration limit. Discipline: Rocket Propulsion The SpaceCraft Object-oriented Rocket Engine Simulation, or SCORES-II, is an SEI COTS product for the analysis of liquid rocket engines6. The tool is written in C++ and features both PC and SGI Unix executable versions. The code, while very similar to the Chemical Equilibrium with Applications (CEA) code in terms of its equilibrium analysis technique and capability, includes a number of added features. These additional features include the ability to analyze throttled engine performance, an “expert-system” database of engine efficiencies to account for engine cycle effects (e.g. staged combustion, expander, gas-generator, etc.), and an engine sizing routine to match a prescribed ambient-pressure thrust value by varying the engine’s propellant mass flowrate. The tool is also capable of modeling almost any generic propellant combination with the user simply providing the propellant composition (e.g. CxHyNz) and initial enthalpy.

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SCORES-II was utilized in the performance estimation and engine sizing for the upperstage’s three HEDM engines. Additionally, SCORES-II is used for modeling the booster and upperstage Reaction Control System (RCS) performance. Discipline: RBCC Propulsion The booster propulsion system analysis was performed using the ‘Simulated Combined Cycle Rocket Engine Analysis Module’ (SCCREAM)7. SCCREAM is a one-dimensional cycle code that is capable of analyzing all modes of an RBCC engines operation. The tool is written in C++, executes on either a PC or SGI Unix machine, and is specifically designed to easily incorporate into the conceptual launch vehicle design environment. Numerous papers are available that document the development of SCCREAM and provide verification test cases with other common industry propulsion tools7. SCCREAM can predict engine performance in a number of different operating modes, including: air-augmented rocket, ramjet, scramjet, scram-rocket, and all-rocket. For the analysis of an engine configuration, the user must define the engine/vehicle geometry, propellant types, component efficiencies, thruster operating conditions, and operating ranges for each mode. The operating ranges for each mode are defined in terms of maximum and minimum altitude, Mach number and throttle setting, with appropriate step increments for each. After analyzing each of the cases, generally at 300-400 different flight conditions, SCCREAM will compile the performance results, in terms of thrust, thrust coefficient and specific impulse, into an engine deck formatted for use with the POST trajectory tool. With this database or map of the engine performance, the trajectory code or analyst has more freedom to explore and determine the optimal flight path. Discipline: Aerodynamics Outer Mold Line (OML) geometries of the booster and upperstage components were created using the Aerodynamic Preliminary Analysis System (APAS) software8. APAS features two different analysis options depending on whether subsonic or supersonic/hypersonic flight conditions are being examined. For subsonic flight, a routine known as

UDP which uses a panel-method analysis approach is invoked. For supersonic and greater flight conditions, a local surface inclination method called Supersonic/Hypersonic Arbitrary-Body Computer Program (S/HABP) is executed. APAS also features a number of engineering approximation methods for estimating skin friction and boundary layer transition points. After the vehicle geometry has been established, an aerodynamic database consisting of lift and drag coefficient tables as functions of Mach number and angle-of-attack are generated across the entire ascent trajectory flight regime. Using a post-processing software script, the APAS results can be compiled into a format that can be read by the trajectory code. At each Mach number and altitude pair of interest, analyses were performed for a range of angles-of-attack. These data tables were provided to the trajectory discipline. Subsequent vehicle scaling was performed photographically and the aerodynamic coefficients were assumed to remain constant during scaling. The aerodynamic analysis was therefore only required at the start of the design process. Note that in the force accounting system used, all forebody pressures were included as aerodynamic drag and the propulsive force was taken to be from the front of the cowl to the tail of the vehicle (cowl-to-tail system). Discipline: Aeroheating For analysis and sizing of the thermal protection systems (TPS), the Thermal Calculations Analysis Tool (TCAT-II) is used. TCAT-II is another SEI COTS tool that solves a one-dimensional unsteady heat transfer equation with an adiabatic backface boundary condition at multiple locations on the vehicle (fuselage, wings, tails, etc.). The tool is written in C++ and features a material property database of over 75 materials including metals, blankets, foams, adhesives, and composites. Designed to read in a vehicle geometry model and convective heating loads from APAS, the number of grid points featured in the APAS model determines the number of individual TPS panels analyzed by TCAT-II. APAS automatically divides the vehicle into various components (body, wing, tail, etc.). For each of these components, the user may specify a unique initial temperature and backface temperature. The flight conditions (time, velocity,

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AOA, etc.) are specified in a simple datafile, and thus can originate from any trajectory code. The user can elect to have TCAT-II either size a TPS stackup by varying the thickness of the materials in the stackup or to simply analyze a specified stackup arrangement. Additionally, the user can specify a list of stackup candidates from which TCAT-II will select the best arrangement at each node. The best arrangement is defined as one that does not exceed its maximum allowable surface temperature limits, meets the maximum backface temperature requirement through either sizing or fixed-thickness analysis, and has the minimum weight. In the event that none of the stackup options meets these criteria, an error warning is generated at the particular node case and the stackup which comes closest to meeting the criteria is selected. Output results from TCAT-II provided to the user include total TPS weight for all components, single component TPS weight, average unit weights by stackup type and component, and coverage or ‘acreage’ percentage by stackup type. Discipline: Operations The Architectural Assessment Tool- Enhanced (AATe) is used in assessing a space transportation system for its operational impacts, mainly costs and cycle times9. It is capable of providing both qualitative and quantitative insights into systems still being conceptualized. The tool is based on the work of both the national Space Propulsion Synergy Team (SPST) and of the joint NASA, Industry & Academia Vision Spaceport project. This model requires both quantitative inputs and qualitative order of magnitude comparisons of the concept vehicle to the Space Shuttle. Inputs include: overall vehicle reliability, airframe life, payload weight, dry weight, vehicle length, and payload demand per year. Outputs include: ground turnaround time, facilities cost, labor cost per flight, line replaceable unit (LRU) cost per flight, and operating expenses per flight. For the Xcalibur concept, a two-stage implementation of AATe was utilized. Discipline: Non-Recurring Cost Non-recurring costs include Design, Development, Testing, and Evaluation (DDT&E) and Theoretical

First Unit (TFU) costs. For this study, an Excel-based model was used that was based on subsystem weight-based Cost Estimating Relationships (CERs) sourced in part from data within NASA’s unrestricted release version of NASCOM database II6. DDTE & TFU was organized around a common WBS similar to that found in the mass properties calculations. Currently, CER’s for the Level 1 WBS are used. The propulsion systems are grouped separately since they are commonly acquired separate from the airframe and associated subsystems. Programmatic costs present in the model include: system test hardware; integration, assembly, & checkout; system test operations; ground support equipment; systems engineering & integration; and program management. Discipline: Reliability and Safety GT-SafetyII is a top-level MS Excel based spreadsheet for determining various safety and reliability metrics for RLVs. The model requires both quantitative inputs from other RLV disciplinary tools as well as specific qualitative user inputs on the architecture being examined (including safety adjustment factors). The quantitative variables consist of variables that describe the physical dimensions of the vehicle (wetted area, length, height, etc.), the configuration (number of propulsion systems, etc.), and the use of the vehicle in the program (flights per year, passengers per flight, etc.). The other types of variables are used for qualitative comparisons of the vehicle in question with the Space Shuttle. The additional safety calculations are separated into the following areas: public/collateral safety, ground personnel safety, flight crew/passenger safety, TPS reliability, engine reliability, and overall mission/vehicle reliability calculations. Output metrics are then determined that relate to both vehicle reliability and program safety. The reliability metrics include terms for both loss of mission and loss of vehicle. The safety metrics include for casualty rate and loss of crew events. Output metrics are listed at the top of a worksheet and are shown in terms of both flights and years between incidents. Discipline: Economics The Cost and Business Analysis Module Abbreviated (CABAM_A) is a MS-Excel spreadsheet-based model that attempts to couple both the demand and supply for space transportation services, specifically RLVs, in the

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future10. The demand takes the form of market assumptions (both near term and far-term) and the supply comes from user-defined vehicles that are placed into the model. CABAM takes inputs from a number of other models to generate Life-Cycle-Cost (LCC) and economic metrics. One of the major assumptions inherent in CABAM is that the project is modeled as a commercial endeavor with the possibility to model the effects of government contribution, tax-breaks, loan guarantees, etc. The model uses 3rd Gen market forecasts based upon the 1994 Commercial Space Transportation Study (CSTS)11.

COLLABORATIVE DESIGN PROCESS The previously described disciplinary engineering models are used in a collaborative engineering framework using Phoenix Integration’s ModelCenter© and Analysis Server© products12. These tools allow the designer to join disparate models and simulations together in a unified environment wherein each discipline can interact with any other discipline. This is performed through a visual interface of an engineering workflow of events where inputs and outputs from various models can be linked together (called a ModelCenter© “model”). This interface allows the engineering process to be more automated and flexible with regards to computing platforms since ModelCenter© and Analysis Server© are relatively platform independent. In addition, these products allow disciplinary models (or ModelCenter© “components”) to be located at diverse geographical locations since data exchange can be performed seamlessly in the environment through the Internet. Driver components besides the models and simulation themselves can be added to the environment. These can include optimizers, trade studies, Design of Experiments (DOE), as well as Monte Carlo components. The interested user should refer to SEI’s own suite of ModelCenter enhancing products, including OptWorks and ProbWorks, for such applications (see the SEI online site for more information, www.sei.aero).

Figure 4. Xcalibur Design Process in ModelCenter© Figure 4 provides a snapshot of the Xcalibur design process implemented in ModelCenter©. The complete vehicle life cycle closure process requires a two step process of closing the performance disciplines and then closing the financial disciplines. Both of these steps are tightly coupled and iterative. For example, a strong iteration loop is present between propulsion, performance (trajectory optimization), and mass properties (weights & sizing). When the vehicle size and capture area changes, the engine performance must be updated and the trajectory re-optimized. During the conceptual design process, the convergence tolerance was taken to be a change of less than 0.1% in gross weight between iterations. The ModelCenter© model setup time is on the order of 20 hours for an experienced user. Execution time for the iterative performance closure process is approximately 3 hours, requiring between 10-15 iterations. The majority of the execution time is devoted to the trajectory optimization using POST. By utilizing this design framework tool, the authors have been able to devote most of the analysis time to refining the disciplinary tools and understanding the design space. Instead of spending hours passing values among disciplines required for the tedious, manual iteration process, the user can simply sit back and monitor the design, devoting more time to understanding how the system reacts to changes. The ModelCenter© model stores all the design and coupling variables, providing an excellent storage device to track the configuration history. Additionally, since the ModelCenter© model is not directly tied to the tools,

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they can be updated over time and the vehicle reconverged by simply re-executing the model.

BASELINE DESIGN RESULTS Aerodynamics The external fuselage configuration of Xcalibur consists of an initial 6.5° half-angle elliptical conic, which transitions to 15° 2-D ramps on the sides of the vehicle and smooth streamlined surfaces on the upper and lower portions. The centerbody was designed to ensure enough volume was available to contain the upperstage. Using APAS the wings were positioned and sized to provide static stability at hypersonic and landing conditions (with flaps). In addition, the wing area was sized to limit the landing speed to fewer than 250 knots. Figures 5 and 6 provide images of the geometric models for the booster that was constructed in APAS.

Figure 5. Booster APAS Model – Top view

Figure 6. Booster APAS Model – Side view For the baseline configuration, the booster required a theoretical wing planform area of 5,790 ft2 (extending into the fuselage). The leading edge sweep of the outboard wing section is 75° and designed so as not to interfere with the engine inlet flowfield. The theoretical aspect ratio of the wing is 0.85 and the taper ratio is 0.08. The wing is a 5% thick biconvex airfoil with a small leading edge radius to reduce wave

drag. The vertical tailfin was sized to have a planform area of 2.5% of the wing theoretical area. The lifting body upperstage/orbiter was designed with a body length-to-width ratio of 2.2. The control surfaces are fully articulating and can be canted downward along the side of the upperstage while stowed in the booster. With the low aerodynamic forces on the upperstage, these surfaces are primarily for control authority during reentry and landing. They provide almost no function during the ascent phase. Internal Configuration & Layout (CAD) The booster for the baseline configuration is approximately 181.7 ft in length (nose-to-tail) and has a total fuselage volume of 66,680 ft3. The maximum width of the booster fuselage is approximately 34 ft. The baseline upperstage vehicle has a length of 42.4 ft, volume of 4,520 ft3, and width of 19 ft. Propellant tanks were packaged in the fuselage of the booster and upperstage vehicles using SDRC-IDEAS, a solid modeling CAD program. Figures 7, 8, 9, and 10 provide both external and internal views of the Xcalibur CAD models used to provide the reference lengths, areas, packaging efficiency calculations, and volumes.

Figure 7. Xcalibur Booster External CAD Image

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Figure 8. Xcalibur Booster Internal CAD Image

Figure 9. Xcalibur Upperstage External CAD Image

Figure 10. Xcalibur Upperstage External CAD Image The Xcalibur booster stage contains provisions for housing three different propellants. The tank volumes for these propellants are sized to provide a total oxidizer to fuel mixture ratio (by weight) of 4.34. A LOX density of 71.2 lbs/ft3, LH2 density of 4.43 lbs/ft3 (normal boiling point), and HEDM density of 52 lbs/ft3 (equivalent to RP) was assumed in both stages.

As shown in Figure 8, the largest tank in the booster is the integral/non-integral LH2 tank, colored in blue. The central body of the booster contains three LOX tanks, two cylindrical and one elliptical, colored in red. The HEDM propellant that is crossfed to the upperstage during takeoff is colored in green and located near the rear of the booster. The packaging efficiency, defined as the ratio of the ascent propellant volume to fuselage volume, is approximately 50% for the booster. This is relatively low and is primarily due to the volume required to house the upperstage vehicle. The main landing gear boxes are located below the two cylindrical LOX tanks. The remaining items present in the booster model are various subsystems including avionics, RCS propellants, Auxiliary Power Units (APUs), and environmental controls. During the design consideration had been given to eliminating the propellant crossfeed and having all upperstage propellants self-contained. The goal would have been to sacrifice upperstage performance for improved operations. The HEDM fuel tank on the booster could have been eliminated as well as the spherical booster LOX tank, reducing system complexity, maintenance, and check-out procedures. However, a detailed study of the above change has not been fully performed. The upperstage lifting body fuselage is dominated by the payload bay volume. The payload bay measured approximately 20’x12’x10’, allowing for a payload density of 8.3 lbs/ft3. Located in front of and behind the payload bay are spherical LOX tanks, colored in red. Behind the payload bay and in the lower left and right sides of the upperstage are cylindrical HEDM fuel tanks. The packaging efficiency for the upperstage is approximately 20%. Propulsion Xcalibur uses four liquid oxygen and hydrogen ejector scram-rocket (ESR) engines on the booster stage to accelerate the upperstage to the Mach 15.5 staging point. As mentioned previously, SCCREAM was used to predict the performance of these engines. Table 2 provides some of the RBCC engine characteristics assumed for the booster.

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Table 2. RBCC Engine Characteristics Item Value

Thrust Chamber Pressure 2,500 psi Thruster Mixture Ratio 6.5 Combustor Efficiency 95.0% Nozzle Efficiency 98.5% Aftbody Expansion Angle 18 degrees

An engine cowl height of 3.8 feet was determined based on a Mach 12 shock-on-lip condition for a conical and 2-D bow shock system. Each engine has an average width of 8.1 feet, allowing for a total inlet area of 30.9 ft2. A variable geometry inlet and exit nozzle were assumed. The total engine length, from cowl leading edge to trailing edge is estimated at 30.5 ft.

0

340,000

680,000

1,020,000

1,360,000

1,700,000

0 50 100 150 200 250 300 350

Time (seconds)

Cow

l-to-

Tai

l Thr

ust (

lbs)

0.0

1.0

2.0

3.0

4.0

5.0

Acceleration (G

's)

ThrustG's

Figure 11. Booster Thrust and Acceleration vs. Time The RBCC thrusters were designed with a chamber pressure of 2,500 lb/in2, and a mixture ratio of 6.5. The excess hydrogen expelled from the thrusters is assumed to react in the constant area mixer section. The engines were sized at sea-level-static (SLS) conditions to meet the vehicle’s overall takeoff thrust-to-weight ratio of 1.15, less the 15% total thrust required contribution from the upperstage. An initial sweep of the system takeoff T/W and %-contribution determined that the 1.15 and 15% values provided the best arrangement for minimizing the total system dry weight. It should be noted that the vehicle system was not extremely sensitive to these settings, within the ranges considered. Each engine is thus capable of producing 244,740 lbs of thrust at SLS, with an Isp of 379 seconds. The engine bypass ratio at SLS was approximately 1.5.

Figure 11 shows the total thrust and vehicle acceleration profile versus time for the booster. Note that the propulsion force accounting system used in SCCREAM is cowl-to-tail. All forebody pressures are included in aerodynamic drag calculated by APAS. Forebody calculations are performed in SCCREAM to determine mass capture at various flight conditions, but the pre-compression effects are not used to reduce the cowl-to-tail thrust values. By design, the booster never switches to the typical RBCC ‘all-rocket’ mode of operation. The scram-rocket mode is capable of operating to the staging point, thus the inlet is never required to be sealed off (minimizing the amount of variable geometry) and the cavity upstream of the thrusters does not require pressurization. Without utilizing a pure-scramjet mode of operation, the SR mode allows for a lower dynamic pressure to be flown from Mach 5.5 to Mach 10. This is in contrast to the scramjet mode operation which requires the vehicle to be flown at the ‘maximum’ dynamic pressure feasible, within available TPS capabilities. All of these factors contribute to reducing the RBCC engine T/W, which was estimated to be 30:1. Performance (Trajectory Optimization) The trajectory for Xcalibur is constrained by changes in pitch rates that provide smooth q-boundary entry and departure paths. The dynamic pressure boundary that is flown is 1,800 lb/ft2 during the ramjet and scram-rocket modes between Mach 2.5 and Mach 10. This provides optimal air-breathing engine performance. The q-boundary is constrained through implementation of a linear feedback control guidance scheme in which the dynamic pressure is held constant by varying the vehicle’s angle-of-attack13. At Mach 10, the vehicle is allowed to depart from the q-boundary and begins to pitch up. Xcalibur flies to a 95 lb/ft2 staging point, where the upperstage is jettisoned. The upperstage then proceeds to inject to a 30x100 nmi. elliptical orbit at a 28.5 degree inclination. The center HEDM engine (1 of 3) functions as the Orbtial Maneuvering System (OMS) propulsion system to circularize the orbit at 100 nmi. and later is used to deorbit the vehicle. The LOX/LH2 OMS system is designed to deliver 450 ft/s of on-orbit ? V.

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Figure 14. Dynamic Pressure vs. Mach Number

Figure 12 shows a graph of the Mach number and altitude versus time results from POST. The angle-of-attack (or alpha) profile prior to staging can be seen in Figure 13. Over the course of the ascent during which the dynamic pressure is held constant (between Mach 2.5 and 10), alpha is fairly constant at 2.5 degrees. A plot of the dynamic pressure as a function of Mach number is given in Figure 14. The 1,800 lb/ft2 q-boundary profile flown during the ramjet mode and a portion of scram-rocket mode can clearly be seen in this figure. Note that with staging occurring at approximately Mach 15.5, the dynamic pressure is 100 lb/ft2. The objective of the trajectory was to minimize the total consumed propellant weight. For the converged baseline (20 Klbs LEO payload), the Mass Ratio (MR) of the ascent was determined to be 3.376 for the booster and 2.158 for the upperstage. The ideal ascent ? V provided by the propulsion system is 31,890 ft/s, including 5,070 ft/s of drag losses (measured inertially). The required total oxidizer-to-fuel mixture ratio was determined to be 4.34 for the booster. The upperstage mixture ratio is always constant at 2.5. Aerothermal Analysis The thermal protection materials and unit weights for the Xcalibur booster were determined using TCAT-II for a representative ascent trajectory and a Mach 3

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flyback cruise. Since the TPS sizing was not performed during every iteration of the vehicle closure process, an offline analysis was required and the TPS weights were updated periodically.

Table 3. Booster TPS Results Item Value

Fuselage TUFI Unit Weight 1.35 lbs/ft2

Acreage 70.2% Fuselage AFRSI

Unit Weight 0.38 lbs/ft2 Acreage 29.8%

Wing TUFI Leeward Unit Weight 0.86 lbs/ft2

Windward Unit Weight 1.65 lbs/ft2 Tail TUFI Unit Weight 1.24 lbs/ft2

For Xcalibur, a TPS design featuring Toughened Uni-Piece Fibrous Insulation (TUFI) tiles, AFRSI blankets, and UHTC’s were selected. Since the exposed wing is constructed of a high-temperature titanium-aluminide, large sections of the wing are designed to be hot structure. Some portions of the wing are covered with TUFI tiles though. To avoid the complexities of active cooling, an ultra-high temperature ceramic (UHTC) is employed on the small radius nosetip and wing leading edges. This material is being developed by NASA Ames and is capable of withstanding temperatures as high as 4,500° F. The UHTC unit weights were estimated at 1.54 lbs/ft.

Figure 15. Booster TCAT-II Results – Top View

Figure 16. Booster TCAT-II Results – Bottom View Table 3 provides a summary of the TPS results generated by TCAT-II for the booster, while Figures 15 and 16 provide the temperature distributions results from the tool. While the unit weights values were fairly close to initial estimates, the high acreage value for the TUFI tiles was not expected (early estimate of ~40%). These acreage values are one of the advantages of the TCAT-II tool and its analyses of the complete vehicle as opposed to a couple of select points. Mass Properties Given a MR (or propellant mass fraction) and a mixture ratio requirement from the trajectory optimization discipline, the model was used to scale the vehicles up (or down) until the available MR matched that required. Once the vehicle was “closed” within the mass properties discipline, the results were sent back to the Propulsion discipline to iterate several times around the Propulsion-Trajectory-Aeroheating-Mass Properties loop. This entire process was repeated, within ModelCenter©, until the gross and dry weights were converged to within 0.1%. The baseline Xcalibur design has a system gross weight of 1,001,500 lbs and a total dry weight of 186,400 lbs. The booster accounts for 162,000 lbs of the dry weight. The upperstage has a gross weight of 103,710 lbs with a dry weight of 24,340 lbs. The fuselage length is 181.7 ft. from tip to tail. Tables 4 and 5 list selected summary items from the Weight Breakdown Structure (WBS) of the booster and upperstage respectively. The full WBS is not included

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in this paper for brevity, but includes 28 major headings with several subcategories under each. Each dry weight component includes a 15% growth margin to account for the likelihood of weight increases.

Table 4. Booster Summary WBS WBS Item Value [lbs]

Body, Wings, and Tails 79,640 TPS 19,610

Landing Gear 5,630 MPS + RCS 46,710 Subsystems 10,470

Dry Weight 162,060 Payload Carried 103,710

Flyback Propellants 21,840 RCS, Unusable, and Reserve Propellants 9,010

Staging Weight 296,620 Ascent Propellants 704,880

GLOW 1,001,500

Table 5. Upperstage Summary WBS

WBS Item Value [lbs] Body and Tails 10,945

TPS 2,290 Landing Gear 1,555

MPS+RCS 4,300 Subsystems 5,250

Dry Weight 24,340 Payload Carried 20,000

RCS, Unusable, and Reserve Propellants 3,730 Insertion Weight 48,070

Ascent Propellants 55,640 GLOW 103,710

Operations Xcalibur is designed to be a highly operable and highly reusable space transportation system. Technologies such as vehicle health monitoring and built-in test equipment are integrated in the design to make checkout and inspection easier, therefore reducing system turnaround time and labor costs. Long life and very reliable airframe (1,000 flights before replacement) and engine components (500 flights before replacement) reduce scheduled maintenance actions and lower inventory costs. The use of toxic fluids such as hypergols has been avoided. The LOX and LH2 propellants are both normal boiling point liquids (no slush LH2). Electro-mechanical actuators are used in place of hydraulics to reduce maintenance costs. The ramjet mode enables the flyback capability and makes it possible to taxi, self-ferry, and obtain improved abort scenarios for Xcalibur.

It is assumed that Xcalibur is operated by a commercial company using a future spaceport and runway at Kennedy Space Center (KSC). The spaceport infrastructure is assumed to be a shared asset provided by the federal or local government similar to today’s airports. Spaceplane operators pay a user’s fee per launch, but are not required to build the spaceport or perform runway maintenance, etc. An estimated streamlined operations crew (approximately 200 “touch” labor personnel for the booster and 150 for the upperstage) are required to operate a single Xcalibur vehicle. The fleet is utilized for an average 25 launches per year (turnaround times of 4 days). The spaceport user’s fee was estimated to be $0.5M (FY2002$) per launch. Economics A conceptual assessment of Xcalibur’s development costs, production costs, fleet size, operational costs, and even potential revenue stream was determined. This assessment was made using SEI’s CABAM_A business model, various SEI and NASA tools, as well as the NAFCOM-99-derived cost model. Within this dynamic economic simulation, changing Xcalibur’s market entry price for each of four different markets (commercial payload, government payload, commercial passenger, government passenger) results in an increase or decrease in potential traffic (if the market is elastic). This price change also affects the number of launches per year required to capture these markets, resulting in a change in annual revenue from each market, and a possible change in fleet size required (due to minimum turnaround time or more often the requirement to replace airframe or engine hardware that has exceed its maximum airframe flight limit). The goal of the optimization was to identify the optimum pricing strategy that results in maximum net present value (NPV), a measure of the financial return of investment projects. For this examination the Xcalibur economic scenario used one market price for two cargo markets (commercial and government) with a fixed passenger flight rate of two and four flights per year for commercial and government passenger markets respectively (at static, fixed prices).

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The Business Environment In this future scenario, Xcalibur will be operated by a private business with the government assisting in the development of the launch service. The U.S. government is a large consumer of launch services and is assumed to be financially interested in lowering launch costs. Therefore, in the following financial data, the government was assumed to pay for 100% of the RBCC engine non-recurring development cost (DDT&E), 100% of the HEDM engine development cost, 25% of the booster and upperstage airframe DDT&E, and 100% of the facilities development. In addition, the government also guaranteed debt loans made to the commercial entity so that financing could be obtained at a reduced interest rate of 7.5%. All airframe and engine production costs as well as all operations and financing costs are borne by the commercial company. The economic environment used in this analysis consisted of an inflation rate of 2.1%, tax rate of 30%, and a discount rate of approximately 15%. A relatively high hurdle rate of 15% is chosen to account for the risky nature of this project. The program starts in 2020 with a projected IOC in 2025 with termination in 2041. A 15% cost margin was added to both DDT&E and Theoretical First Unit (TFU) costs (this was use din place of the weight margin). Economic Results The optimized business scenario is shown here for a Weighted Average Cost of Capital (WACC)+ Incentive Return (IR) of 15.23% with a fleet size of only one vehicle and a total steady state flight rate of 25 launches per year (40.7% commercial cargo). Incentive Return is an additive parameter to the WACC that SEI occasionally uses in RLV financial analysis. This percentage is added to the WACC to create a new discount rate. This is an extra incentive rate required, beyond the NPV just being positive, for the project to be accepted. The launch vehicle company operates for 21 steady state years after a two year ramp up and launches a total of approximately 500 times. The venture is predicted to incur a total Life Cycle Cost (LCC) of $23.24B (FY2002$) with an initial debt-to-equity ratio of 3. Non-recurring costs (DDT&E, engine and airframe production, but not financing costs) of the entire venture are estimated to

be $9.73B (FY2002$) of which the U.S. government is expected to contribute $6B (FY2002$) over the life of the project (DDT&E, facilities, etc.). The total cost incurred from program start to initial vehicle acquisition is $12.431B which includes the total DDT&E, airframe TFU’s, and all engines to complete a single vehicle system. Table 6 contains a more detailed summary of the non-recurring cost results. The optimized market price is equivalent to current launch prices on an expendable system. The reductions are more significant with respect to the Space Shuttle. However, dramatic (orders of magnitude) decreases in access to space costs do not appear likely given the current models and assumptions if the proposed company is to achieve an attractive rate of return given investment risk.

Table 6. Xcalibur Economic Results Summary Non-Recurring Cost Component Booster Airframe Total DDT&E

Airframe $5.85B $1.92B $7.77B Engines $7.72B $2.02B $9.74B

TFU Airframe $1.24B $0.39B $1.63B Engines $0.30B $0.03B $0.32B

Total Engine Acquisition $1.00B $0.06B $1.06B First Vehicle Acquisition $2.24B $0.45B $2.70B Total Cost First Vehicle $12.43B Note: All figures in FY2002$, with learning curve effects of 85% on engines For Xcalibur, fairly aggressive assumptions were made when determining recurring costs. For this study, recurring costs were assumed to be the sum of the following four items (in FY2002 dollars): 1.) labor costs, 2.) Line Replacement Unit (LRU) costs, 3.) Propellant costs at $0.10/lb of LOX and $0.25/lb of LH2 based on the assumption of an on-site propellant production facility, and 4.) Insurance costs of approximately $50,000/launch based as a function of Loss of Vehicle (LOV) reliability and a 5% insurance premium. This insurance is only for limited liability coverage for the vehicle (hull insurance) and does not include payload coverage and liability for public collateral injury (assumptions include government indemnity for such occurrences). Based on these assumptions, each flight of Xcalibur is estimated to cost $8.3M (FY2002$). The recurring cost per pound for delivering payloads (20,000 lb) to a 100 nmi. due east orbit is $416/lb. This cost includes operations, propellant, insurance, and the site fee. The

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direct recurring operations cost includes only operations and propellant costs. It is important to note that these recurring costs do not convey the actual market value for these launch services. Actual Xcalibur customers pay the optimized launch price in Table 7, not the recurring cost. The price includes recurring costs, amortized hardware and design costs, financing costs, and company profit and thus is several times higher.

Table 7. Xcalibur Economic Results Summary Recurring Cost Component Value

Market Price per pound $5,542 Recurring Cost per pound $416 Direct Recurring Cost per pound $307 Total LCC (w/o financing, pre-govt contr.) $23.34B Total Government Contribution $6.06B Note: All figures in FY2002$

SUBSCALE DEMONSTRATOR VEHICLE A subsequent examination was performed to estimate the size and cost of developing a prototype version of the booster. The upperstage vehicle component was completely eliminated from the system, by effectively reducing the booster payload to zero. The internal volume fraction for the upperstage was maintained though. The RBCC engine T/W was reduced from 30:1 to 20:1, but the performance in terms of thrust and Isp was assumed to be achievable. The SLS thrust of the RBCC engines were also increased to make up for the 15% thrust contribution provided by the upperstage in the full-scale vehicle. The airframe structural and tank unit weights were also increased, but all subsystem weights (landing gear, TPS, EMAs, etc.) assumed similar advanced technology levels as the full-scale system. Additionally, the vehicle was only required to accelerate up to Mach 12, thus reducing the propellant load and flyback distance. Operation of the prototype vehicle was not considered, only the DDT&E and TFU required in building the prototype system. Table 8 provides a summary of the prototype vehicle properties.

Table 8. Prototype Xcalibur Specifications Booster Values

Length 104.1 ft Dry Weight 49,470 lbs. GLOW 176,680 lbs. DDT&E $2.80B Acquisition $1.09B Total Cost to Acquire $3.89B Note: All figures in FY2002$

While the GLOW and dry weight of the prototype system appears to be very attractive, there is still a large investment hurdle of almost $4B required to obtain the system. This is primarily due to the required technology developments. Considering the high level of uncertainty and experience from past RLV development efforts (e.g. X-33), this value is most likely a conservative estimate. Substantial cost-sharing (>80%) by the government would most likely be required to initiate this project.

PAYLOAD TRADE STUDY

In an effort to reduce the DDT&E investment hurdle, an alternate payload configuration was examined. It was hoped that the smaller vehicle would offer a significant reduction in vehicle size and development cost, thus allowing a lower price-per-pound to be charged to the customer resulting in turn increase the flight rates and vehicle utilization. The increase in flight rates should also help to improve vehicle operations and reduce the recurring vehicle costs. The payload weight was reduced from the baseline 20 Klbs to 10 Klbs, to LEO at 28.5o. The payload bay volume was also reduced by approximately 75%, greatly improving the packaging efficiency in the upperstage and reducing the vehicle size. The smaller upperstage also allowed for a smaller booster vehicle. After closing the performance disciplines for this configuration, a system GLOW of 570,100 lbs and a total dry weight of 113,100 lbs was obtained. The booster length was reduced from 181.7 ft to 148.9 ft, and the upperstage was reduced to 28.4 ft in length. A slight improvement in vehicle mass ratio was also obtained due to scaling effects. Table 9 provides a summary of the results.

Table 9. Payload Trade Study Results for 10 Klbs Payload Vehicle

Item Values Booster Length 148.9 ft Booster Dry Weight 98,900 lbs. Booster Stage Weight 514,700 lbs. Upperstage Length 28.4 ft Upperstage Dry Weight 14,200 lbs. Upperstage Stage Weight 55,400 lbs. DDT&E $7.55B Price per pound $4,580/lb. Flight Rate 55 per year Recurring Cost per flight $6.4M Note: All figures in FY2002$

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SUMMARY

SpaceWorks Engineering, Inc. (SEI) was tasked by NASA MSFC’s Advanced Concepts group to investigate the feasibility and viability of a two-stage, vertical takeoff with horizontal landing system utilizing RBCC propulsion. This request resulted in the design and full life cycle analysis of the Xcalibur concept by SEI. The design of the Xcalibur system was accomplished in a distributive multidisciplinary, multi-code automated environment. The framework utilized to support this process was Phoenix Integration’s ModelCenter© collaborative design environment. Given the complexity of the analysis presented here, this environment proved highly valuable in the design process. The Xcalibur concept appears to offer a feasible design as a 3rd Generation two-stage, horizontal takeoff and vertical landing RLV. A fairly attractive system GLOW weight of just over 1 Mlbs was obtained for delivery of 20 Klbs of payload at orbit of 28.5o LEO. The dry weight of the booster stage was determined to be 162,060 lbs and the upperstage was 24,340 lbs. Extended operation of the booster engine’s RBCC scram-rocket mode was utilized to significantly improve vehicle acceleration and reduce the vehicle engine/airframe sensitivity necessary for high-Mach scramjet operation. The two-stage configuration and staging point of approximately Mach 15.5 at a dynamic pressure of 100 lb/ft2 was selected to eliminate the need for transitioning to the RBCC engines all-rocket mode. The performance in this mode typically has a high degree of uncertainty and its elimination reduces concept risk. Additionally, a higher engine T/W is achievable by eliminating the need to close and seal off the inlet doors (less variable geometry) that also reduced the requirement for pressurization gas to fill the resultant diffuser/isolator cavity. The HEDM engines of the upperstage system provided excellent performance and resulted in a very small, dense vehicle. While it was not demonstrated in this brief, this technology can be labeled as an “enhancing”

(versus an “enabling” technology). Thus, replacing the HEDM propulsion system with advanced kerosene engines will result in some vehicle growth. This technology replacement may need more investigation. The economic results of the Xcalibur system are generally somewhat disappointing. While an aggressive commercial-run (with some government backing) scenario was attempted, the vehicle still failed to obtain a viable economic case (Net Present Value>$0) for a discount rate of approximately 15%. The DDT&E was $9.74B for the 20 Klbs and $7.55B for the 10 Klbs payload configurations, but this large upfront cost drove the market prices to values on par with current expendable launch vehicles at $4K-$5K/lb.

ACKNOWLEDGEMENTS

The authors wish to acknowledge the National Aeronautics and Space Administration (NASA) and the Advanced Concepts group at Marshall Space Flight Center (MSFC) for sponsorship of this work by SpaceWorks Engineering, Inc. (SEI). The authors would also like to acknowledge Phoenix Integration, Inc. for providing assistance in the development of the collaborative environment with access to the latest ModelCenter© and Analysis Server© products.

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