NASA Space Launch System Operations Strategy

The National Aeronautics and Space Administration (NASA) is developing new capabilities for human and scientific exploration beyond Earth's orbit. This effort includes the Space Shuttle derived Space Launch System (SLS), the Multi-Purpose Crew Vehicle (MPCV) and the Ground Systems Development and Operations (GSDO). There are several requirements and Technical Performance Measures (TPMs) that have been levied by the Exploration Systems Development (ESD) upon the SLS, MPCV, and GSDO Programs including an integrated Launch Availability (LA) TPM. The LA TPM is used to drive into the SLS, Orion and GSDO designs a high confidence of successfully launching exploration missions that have narrow Earth departure windows. The LA TPM takes into consideration the reliability of the overall system (SLS, Orion and GSDO), natural environments, likelihood of a failure, and the time required to recover from an anomaly. A challenge with the LA TPM is the interrelationships between SLS, Orion, GSDO and the natural environments during launch countdown and launch delays that makes it impossible to develop an analytical solution for calculating the integrated launch probability. This paper provides an overview of how Discrete Event Simulation (DES) modeling was used to develop the LA TPM, how it was allocated down to the individual programs, and how the LA analysis is being used to inform and drive the SLS, Orion, and GSDO designs to ensure adequate launch availability for future human exploration.

Exciting and robust human and robotic exploration missions will be enabled by the NASA Space Launch System (SLS). The SLS's capabilities for launching heavy payloads with large diameters (up to 10 meters) will enable a variety of science rich missions including Mars and outer planet missions; including Human missions utilizing the Orion Crew Vehicle. Missions to cislunar destinations will serve as proving grounds for later Crewed Mars Missions in the 2030s. In this report an outline is given of the capabilities of the SLS, strategies for exploration missions are presented, and descriptions of SLS based payload elements are given. The SLS program is making progress toward delivering a new capability for Beyond Earth Orbit exploration.

NASA's Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher characteristic energy (C3) energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as “monolithic” telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

The National Aeronautics and Space Administration has formed the Advanced Space Transportation Program (ASTP) at the Marshall Space Flight Center to address advanced space transportation technologies. The main focus of the ASTP Office is directed to those technologies that will be needed to reduce the cost of both Earth-toorbit and in-space transportation. The current focus of the advanced reusable technologies (ART) Project is the development of those critical technologies required to enable a rocket-based combined-cycle (RBCC) engine and ultimately an RBCC-based reusable launch vehicle. Additional effort is also being expended to address those associated technologies that would be required to not only support an RBCC-based vehicle but other advanced reusable transportation systems as well. This paper will describe the work that has been performed since the last project status was presented to the 1997 Joint Propulsion Conference.' Currently, NASA and its industry partners are performing ground testing of hydrogen-fueled rocketbased combined-cycle flowpaths. Successful ramjet and scramjet testing at Mach 6 and scramjet testing at Mach 8 have been performed. Cold flow mixing tests have also been successfully performed as have inlet operability tests. Additional testing of RBCC flowpaths at air augmented rocket and rocket only modes is underway. Additional work will be performed on technologies that will support an RBCC-powered launch vehicle and will be focused in the following areas: structures and materials; avionics and operations; propulsion turbomachinery; and thermal protection systems. Chief Engineer, Advanced Reusable Technologies Project, NASA/ MSFC, Senior Member AIAA Project Manager, Advanced Reusable Technologies Project, NASA/MSFC, Member AIAA Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental Purposes. All other rights are reserved by the copyright owner. Introduction The National Aeronautics and Space Administration (NASA) Advanced Space Transportation Program (ASTP), part of the Space Transportation Program (STP) Office at the Marshall Space Flight Center (MSFC), is focusing on those technologies that will be needed to reduce the cost of both Earth-to-orbit and in-space transportation. NASA's strategy is to advance innovative space transportation technology development through the X-33, X-34, ASTP, and the Future-X programs to enable reduction of these costs. The ASTP consists of three major technology areas: focused, core, and research technologies. The ART project currently manages the work that is covered under NASA Research Announcements 8-16 and 8-21. The project's primary focus over the last 2 years has been on rocket-based combined-cycle (RBCC) technology development activities awarded under NRA 8-16. The current aerospace and academic organizations listed in Figure 1 are performing testing and analysis supporting the RBCC propulsion activities. The engine contractors, Aerojet, Pratt & Whitney, and Rocketdyne are currently either testing or preparing to test their RBCC engine flowpaths. Direct connect testing has been performed at Mach 6 flight conditions in both ramjet and scramjet modes and at Mach 8 flight conditions in the scramjet mode. Testing has also been performed at sealevel static conditions and at Mach 3.4 ramjet conditions.

This paper provides an overview of the tools used to conduct independent modeling and simulation of Artemis 1 separation events and highlights challenges and lessons learned across a decade of analysis. An overview of each of the tools along with the process flow and key components of maintaining independence from the Program’s baseline tool set will be provided. In addition, each of the separation analyses conducted will be discussed, making note of key models incorporated for each event and unique challenges that were addressed. The technical work described in this paper is part of an on-going technical assessment conducted by the NASA Engineering and Safety Center to develop and maintain an independent modeling and simulation capability in support of Space Launch System (SLS) and Multi-Purpose Crew Vehicle (MPCV) throughout design, verification, and flight readiness cycles. For Artemis 1, an end-to-end analysis approach was taken with events simulated from launch to re-entry. This paper will focus on the separation events, including liftoff clearance analysis, solid rocket booster (SRB) separation from Core Stage (CS), Service Module (SM) panel jettison, ICPS / Core Stage Separation, and MPCV/ ICPS separation. To assess separation clearances, the NESC has developed a high fidelity, independent 6-degree of freedom (DOF) simulation of the separation events. The simulation was developed using the Program to Optimize Simulated Trajectories II (POST2). POST2 is an industry standard flight trajectory optimization and simulation tool that is Class D compliant and follows NASA NPR 7150.2, which includes regression testing, unit testing, and configuration control of the simulation. The primary objective of the SLS separation and clearance analyses was to estimate the clearances between different vehicle components and/or ground support equipment (GSE) during separation events. The clearance is defined as the minimum distance from any point on one 3D geometry model associated with one vehicle component (or GSE) to any point on another 3D geometry model associated with a vehicle component separating from the main body. The vehicle geometry models used by the NESC to compute the separation clearances were derived from CAD models. POST2 simulation results were post processed in the Exploration Visualization Environment (EVE) tool to compute clearances, and detect recontact occurrences, if any, between the separating stages. EVE is a simulation, visualization, and analysis system designed to integrate time-based dynamics data with detailed graphical models in a full-scale virtual environment. EVE was used to simultaneously calculate more than 30 different separation clearances by “driving” detailed vehicle geometry models with position and attitude time history data obtained from the POST2 simulation. EVE displays each clearance distance visually, at any given instant of time, with a straight line connecting the two closest points on each object and records the time history of the clearance distance in a file. If the clearance reduces to zero, then EVE reports it as a recontact. To perform separation analyses, POST2 Monte Carlo analyses were performed by running 2000 dispersed cases that included over hundreds of uncertainties and dispersions for a range of simulation models including aerodynamics, propulsion, navigation sensors, control actuators, slosh and flexible body dynamics, and winds and atmosphere. Clearances were computed for each of the dispersed Monte Carlo trajectories and overall separation performance was statistically assessed. Over the past decade the POST2 simulation and EVE visualization tools have been used to perform detailed assessments of the separation events that occur during the Artemis 1 ascent. To support these assessments, numerous unique simulation models were developed and integrated into the standard suite of POST2 simulation models to address and resolve key concerns specific to each of the Artemis 1 separation events. In addition to presenting an overview of the multi-body 6-DOF simulation and separation analysis capability, the paper will discuss the integration of key simulation models such as flexible body dynamics, slosh dynamics, separation mechanisms, multi-body aerodynamics, and environmental models. Additional details are provided about other challenges that were addressed, including the use of a convex hull to reduce computational time for clearance calculations, the modeling of incidental recontact during separation events, and the capability to simulate failure scenarios.

We present a design for an active telescope for space astronomy. The telescope is capable of both exoplanet work and general astronomy over wavelengths from ∼100 nm up to 5 μm. The primary mirror is 6 m in diameter, formed by 16 mirror segments that are precisely phased and supported on rigid body actuators and with segment optical surface figures fine-tuned using surface figure actuators. The active primary forms a large deformable mirror (DM) with wavefront error (WFE) correction at the entrance pupil. Thus the largest source of WFE can be removed at the source and is corrected over the entire field of view. This enables diffraction-limited performance at 400 nm and a more efficient optical system over a broader wavelength range than could be achieved by a small DM at a downstream relayed pupil. The telescope is passively cooled to below 100 K at Sun–Earth L2, enabling astronomical-background-limited observations out to 5 μm. Launched on a SpaceX Starship or alternatively National Aeronautics and Space Administration’s Space Launch System, the telescope requires minimal deployments. A 72-m-diameter starshade provides a contrast ratio better than 10 − 10 for exoplanet science. Near the visible region, with a 108% working bandwidth from 300 to 1000 nm, a working distance of 120 Mm provides a 51-mas inner working angle (IWA). This band can be moved to shorter or longer wavelengths by adjusting the starshade range from the telescope. Our first-ever thermal analysis of such a starshade shows that a temperature below 100 K can be achieved over a broad range of observing directions, permitting the possibility of working into the infrared. We model the yield in exoplanets that can be observed. A starshade and associated spectrograph offer significant advantages for exoplanet characterization. They enable a much broader instantaneous spectral bandwidth (here 108%) than current coronagraphs (∼10 % to 20% bandwidth), allow both polarizations to be observed simultaneously, and have higher throughput. The IWA is twice as small as can be achieved with a coronagraph and there is no outer working angle. These differences are particularly pronounced in the UV, where coronagraph performance would be strongly affected by throughput losses, wavefront aberrations, Fresnel polarization effects at surfaces, and thermal instability.

This manuscript presents experimental results of compression tests of repaired sandwich structure typical of a launch vehicle Payload Adapter Fitting (PAF) that is being considered for the National Aeronautics and Space Administration’s (NASA) Space Launch System (SLS) launch vehicle. Specimens with 8-ply quasi-isotropic carbon/epoxy face sheets and aluminum honeycomb core were damaged with representations of misdrilled holes, barely visible impact damage (BVID) or visible impact damage (VID). A patch repair technique not requiring removal of any material is presented and assessed for residual in-plane compressive load-carrying capability on each of the three types of damage. Edgewise compression tests on specimens demonstrated the repairs could recover all measured undamaged compressive load-carrying capability of the misdrilled hole and BVID specimens. For the specimens with VID, an average of about 91% of the average measured undamaged compressive load carrying capability was recovered indicating that larger patches, or perhaps removal and replacement of the damaged material, may be needed for more severe damage if all the measured undamaged compression strength needs to be regained.

Solar sailing is a unique form of propulsion where a spacecraft gains momentum from incident photons. Solar sails are not limited by reaction mass and provide continual acceleration, reduced only by the lifetime of the lightweight film in the space environment and the distance to the Sun. Once thought to be difficult or impossible, solar sailing has come out of science fiction and into the realm of possibility. Any spacecraft using this method would need to deploy a thin sail that could be as large as many kilometers in extent. The availability of strong, ultra lightweight, and radiation resistant materials will determine the future of solar sailing. The National Aeronautics and Space Administration's Marshall Space Flight Center (MSFC) is concentrating research into the utilization of ultra lightweight materials for spacecraft propulsion. The Space Environmental Effects Team at MSFC is actively characterizing candidate solar sail material to evaluate the thermo-optical and mechanical properties after exposure to space environmental effects. This paper will describe the exposure of candidate solar sail materials to emulated space environmental effects including energetic electrons, combined electrons and Ultraviolet radiation, and hypervelocity impact of irradiated solar sail material. This paper will describe the testing procedure and the material characterization results of this investigation.

We have been studying an architecture for Cislunar Development that includes early deployment of an Exploration Platform at one of the Earth – Moon Lagrange points. The Exploration Platform provides a flexible basis for future exploration, since it reduces cost through reuse of expensive vehicles and reduces the number of launches needed to accomplish missions. International Space Station (ISS) industry partners have been working for the past two years on concepts for using ISS development methods and residual assets to support a broad range of exploration missions. These concepts have matured along with planning details for NASA’s Space Launch System (SLS) and Multi-Purpose Crew Vehicle (MPCV) to allow serious consideration for a platform located in the Earth-Moon Libration (EML) system. We propose using the mass margin available on the 2017 uncrewed ORION/SLS test flight to launch secondary science payloads that advance science and exploration objectives.

Interstellar Probe: A Mission to Explore the Heliospheric Boundary and Interstellar Medium

Personal computers of the past were limited to running computational fluid dynamics codes in serial mode. With the advent of multicore technology and suitable compilers, personal computers nowadays can execute codes in a parallel fashion similar to that of supercomputers and cluster computer systems. This article investigated what performance can be achieved when executing an aerodynamic code on an Intel quad core based personal computer with an OpenMP compiler in a Windows environment. The code solved the Euler equations to find the flowfield around a NACA0012 aerofoil on an O-type boundary fitted structured grid system. A speedup of up to 350% was obtainable in double precision accuracy. The reduced computation time means that, for small scale problems, more accurate Euler codes can replace commonly used transonic small disturbance codes. For larger problems, this information serves as a reference for developers of hybrid MPI/OpenMP algorithms for cluster computer systems. References Advanced Micro Devices (AMD), AMD FireStream 9170: Industry's First GPU with Double-Precision Floating Point, 2007, http://ati.amd.com/products/streamprocessor/specs.html Allmaras, S. R., Venkatakrishnan, V., and Johnson, F. T., Farfield Boundary Conditions for 2-D Airfoils, AIAA Paper 2005-4711, AIAA, 2005. Bailey, D., Barszcz, E., Barton, J., Browning, D., Carter, R., Dagum, L., Fatoohi, R., Fineberg, S., Frederickson, P., Lasinski, T., Schreiber, R., Simon, H., Venkatakrishnan, V., and Weeratunga, S., The NAS Parallel Benchmarks, National Aeronautics and Space Administration (NASA), Contractor Report 203186, USA, 1994, http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19970014936_1997021577.pdf Dong, S., and Karniadakis, G. E., Dual-Level Parallelism for High-Order CFD Methods, Parallel Computing, 30, Jan. 2004, pp. 1--20. Hirsch, C., Numerical Computation of Internal and External Flows (Volume 2): Computational Methods for Inviscid and Viscous Flows, John Wiley and Sons, Great Britain, 1992. Hoffmann, K. A., Computational Fluid Dynamics for Engineers, Engineering Educational System, Texas, USA, 1989. Intel, Tera-scale Computing Research Program, c.2007, http://www.intel.com/research/platform/terascale/ Jin, H., Frumkin, M., and Yan, J., The OpenMP Implementation of NAS Parallel Benchmarks and its Performance, National Aeronautics and Space Administration (NASA), Technical Report NAS-99-011, Moffett Field, USA, 1999. Liu, Y., and Vinokur, M., Nonequilibrium Flow Computations 1: An Analysis of Numerical Formulations of Conservation Laws, National Aeronautics and Space Administration (NASA), Contractor Report 177489, California, USA, 1988, http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19880021000_1988021000.pdf Lock, R. C., Test Cases for Numerical Methods in Two-Dimensional Transonic Flows, Advisory Group for Aerospace Research and Development (AGARD), Report Number 575, 1970. Ly, E., Improved Approximate Factorisation Algorithm for the Steady Subsonic and Transonic Flow over an Aircraft Wing, Proceedings of the 21st Congress of the International Council of the Aeronautical Sciences (ICAS98), ICAS and AIAA, Melbourne, Australia, 1998, Paper A98-31699. Ly, E., and Gear, J. A., Time-Linearized Transonic Computations Including Shock Wave Motion Effects, Journal of Aircraft, 39, 6, Nov/Dec. 2002, pp. 964--972. Ly, E., and Nakamichi, J., Time-Linearised Transonic Computations Including Entropy, Vorticity and Shock Wave Motion Effects, The Aeronautical Journal, Nov. 2003, pp. 687--695. 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A successful transition to global clean energy hinges on meeting the world’s growing demand for power, while at the same time reducing greenhouse gas emissions. Achieving this will require significant growth in electricity generation from clean and carbon-free energy sources. Several energy providers (Acar and Dincer, 2019, “Review and Evaluation of Hydrogen Production Options for Better Environment,” J. Cleaner Prod., 218, pp. 835–849 and Nagashima, M., 2024, “Japan’s Hydrogen Strategy and Its Economic and Geopolitical Implications,” Afficher la page d’accueil du site) have already begun the transition from traditional carbon-based fuels to cleaner alternatives, such as hydrogen (H2) and hydrogen-enriched natural gas (HENG). However, there are still many technical questions/challenges that must be addressed when applying these fuels in gas turbines. The application of H2 or H2/natural gas (NG) blends to advanced-class gas turbines, which have higher operating pressures and temperatures, has raised concerns about the potential for leakages or fuel sequencing operations where flammable mixtures of fuel and air could auto-ignite. Public information on the auto-ignition of H2 in the air at atmospheric pressure shows an auto-ignition temperature (AIT) between 520 and 585 °C (U.S. Department of Energy, 2024, “U.S. Department of Energy Hydrogen Program Plan,” U.S. Department of Energy, Washington, DC; Georgia Power, 2022, “Georgia Power, Mitsubishi Power, EPRI Complete World’s Largest Hydrogen Fuel Blending at Plant McDonough-Atkinson,” Georgia Power, Atlanta, GA; and gepower-v2, 2024, “New York Power Authority: GE Vernova,” gepower-v2, Cambridge, MA). Such data show AIT of H2 is ∼100 °C lower than that of methane (CH4) which has a minimum AIT of around 600 °C (Huth, M., and Heilos, A., 2013, “Fuel Flexibility in Gas Turbine Systems: Impact on Burner Design and Performance,” Modern Gas Turbine Systems, Sawston, UK, pp. 635–684). Studies also show that as pressure increases, methane’s AIT decreases significantly to around 390 °C (Loving, C., Mastantuono, G., Terracciano, A. C., Vasu, S. S., Pigon, T., Hernandez, A., and Cloyd, S., 2023, “Auto-Ignition Test Results of Hydrogen and Natural Gas Fuels at Atmospheric and Elevated Pressures for Gas Turbine Safety,” ASME Paper No. GT2023-102674). However, there was insufficient information in the published literature to characterize the influence of pressure on the AIT of H2 and HENG fuels. At atmospheric conditions, H2 has a wider flammability range of equivalence ratios that ignition can occur compared to methane. H2’s flammability ranges from 4% to 75% molar (volume) fuel concentration, which is an equivalence ratio range of 0.137–2.57. Methane’s flammability limit ranges from 5% to 15% molar (volume) or an equivalence ratio between 0.53 and 1.58 (National Aeronautics and Space Administration, 1997, Safety Standard for Hydrogen and Hydrogen Systems: Guidelines for Hydrogen System Design, Materials Selection, Operations, Storage, and Transportation, National Aeronautics and Space Administration, Office of Safety and Mission Assurance; National Technical Information Service, Distributor, Washington, DC, Springfield, VA). Previous research has also been done to determine the effect of longer hydrocarbons present in natural gas mixtures. The presence of ethane (C2H6) and propane (C3H8) has been shown to reduce the AIT of natural gas, especially at elevated pressures (The Association, 1994, NFPA 49: Hazardous Chemicals Data, The Association, Quincy, MA). These longer hydrocarbons also tend to promote ignition in richer conditions, whereas methane tends to ignite easier in slightly lean conditions (National Aeronautics and Space Administration, 1997, Safety Standard for Hydrogen and Hydrogen Systems: Guidelines for Hydrogen System Design, Materials Selection, Operations, Storage, and Transportation, National Aeronautics and Space Administration, Office of Safety and Mission Assurance; National Technical Information Service, Distributor, Washington, DC, Springfield, VA). Numerous variables besides the pressure, fuel, and equivalence ratio can affect the AIT including chamber volume size, chamber materials, presence of diluents, and other factors (Standards Australia International, 2000, Electrical Apparatus for Explosive Gas Atmospheres. Part 20, Data for Flammable Gases and Vapours, Relating to the Use of Electrical Apparatus, Standards Australia International; Standards New Zealand, Strathfield, NSW, Wellington, NZ). This study describes the test methodology used to evaluate conditions where auto-ignition occurs for various fuel–air mixtures operating at different pressures (1–30 atm) and temperatures. Testing was completed with 100% H2 and multiple H2/NG blends at various equivalence ratios (ER) between 0.2 and 2.5. Testing was similarly performed for 100% NG to validate the test and data collection methods cited in prior published literature. Results indicate that, at atmospheric pressures, an increase in H2 concentration results in a reduced AIT. However, at 30 atm, the increased presence of H2 increased the AIT. At elevated pressures above 10 atm, increased equivalence ratio resulted in reduced AIT for all mixtures with NG having the greatest sensitivity to equivalence ratio. Variations of auto-ignition delay times (AIDT) were also observed during the testing and are compared to modeling predictions, providing insight into auto-ignition characteristics.

This paper describes the results of a study to determine the performance improvements achievable by circumferentially indexing successive rows of turbine stator airfoils. An experimental / analytical investigation has been completed which indicates significant stage efficiency increases can be attained through application of this airfoil clocking concept. A series of tests was conducted at the National Aeronautics and Space Administration’s (NASA) Marshall Space Flight Center (MSFC) to experimentally investigate stator wake clocking effects on the performance of the Space Shuttle Main Engine Alternate Fuel Turbopump Turbine Test Article. Extensive time-accurate Computational Fluid Dynamics (CFD) simulations have been completed for the test configurations. The CFD results provide insight into the performance improvement mechanism. Part one of this paper describes details of the test facility, rig geometry, instrumentation, and aerodynamic operating parameters. Results of turbine testing at the aerodynamic design point are presented for six circumferential positions of the first stage stator, along with a description of the initial CFD analyses performed for the test article. It should be noted that first vane positions 1 and 6 produced identical first to second vane indexing. Results obtained from off-design testing of the “best” and “worst” stator clocking positions, and testing over a range of Reynolds numbers are also presented. Part two of this paper describes the numerical simulations performed in support of the experimental test program described in part one. Time-accurate Navier-Stokes flow analyses have been completed for the five different turbine stator positions tested. Details of the computational procedure and results are presented. Analysis results include predictions of instantaneous and time-average mid-span airfoil and turbine performance, as well as gas conditions throughout the flow field. An initial understanding of the turbine performance improvement mechanism is described.

The Transient Pressure Test Article (TPTA) test program is being conducted at a new test facility located in the east test area at the National Aeronautics and Space Administration's (NASA's) Marshall Space Flight Center (MSFC) in Huntsville, Alabama. This facility, along with the special test equipment (STE) required for facility support, was constructed specifically to test and verify the sealing capability of the redesigned solid rocket motor's (RSRM) field, igniter, and nozzle joints. The test article consists of full scale RSRM hardware loaded with inert propellant and assembled in a short stack configuration. The TPTA is pressurized by igniting a propellant cartridge capable of inducing a pressure rise rate that simulates the ignition transient that occurs during launch. Dynamic loads are applied during the pressure cycle to simulate external tank attach (ETA) strut loads present on the ETA ring. Sealing ability of the redesigned joints is evaluated under joint movement conditions produced by these combined loads since joint sealing ability depends on seal resilience velocity being greater than gap opening velocity. Also, maximum flight dynamic loads are applied to the test article which is either pressurized to 600 psia using gaseous nitrogen (GN2) or applied to the test article as the pressure decays inside the test article on the down cycle after the ignition transient cycle. This paper will present the uniqueness of this new test facility with respect to its capabilities. It will also touch on both the topic of test effectiveness versus space vehicle flight performance and new aerospace test techniques, as well as presenting a comparison between the old SRM design and the RSRM.

Although the Orion Multi-Purpose Crew Vehicle (MPCV) is being designed primarily for water landings, the possibility exists of an onshore landing at Kennedy Space Center (KSC) following a launch abort. To obtain data to evaluate the adequacy of sand material models used in beach landing simulations, sand impact tests were conducted at National Aeronautics and Space Administration (NASA) Langley Research Center (LRC). Both vertical drop tests and swing tests with combined vertical and horizontal velocity were performed onto beds of common construction-grade sand using a geometrically scaled crew module test article. The tests were simulated using the explicit, nonlinear, transient dynamic finite-element code LS-DYNA. The shape of the test article was similar to a scaled crew module, but the structure was much heavier and was sufficiently robust to survive multiple impacts without permanent deformation. The focus was the sand model, not the structural response of the test article. The sand material ...