Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Supersonic Retropropulsion”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

An Overview of NASA's Current Materials Development Efforts for Mars EDL

Current roadmaps point to landing heavy masses (cargo, followed by manned vehicles) on Mars in the 2030's and the existing entry, descent and landing (EDL) technology will not be sufficient to facilitate such missions. In 2009 the Exploration Technology Development Program (ETDP) established the Entry, Descent and Landing Technology Development Project (EDL TDP), to be managed programmatically at Langley Research Center (LaRC) and technically a Ames Research Center (ARC). The purpose of the project is to further the technologies required to land heavy (approximately 40 metric ton) masses on Mars to facilitate exploration. The EDL TDP contains three technical elements. They are: 1) Thermal Protection Systems (TPS) development 2) Modeling and Tools (MAT) development 3) Supersonic Retropropulsion (SRP) development The primary goals of the EDL TDP TPS element is to design and develop TPS materials capable of withstanding the severe aerothermal loads associated with aerocapture and entry into the Martian atmosphere while significantly decreasing the TPS mass fraction contribution to the entry system. Significant advancements in TPS materials technology are needed in order to enable heavy mass payloads to be successfully landed on the Martian surface for robotic precursors and subsequent human exploration missions. The EDL TDP TPS element is further divided into two different TPS concepts for Mars EDL those being: 1) Rigid TPS for a mid L/D aeroshell with the capability to withstand dual pulsed heating environments as high as 500 W/square cm for aerocapture and 130 W/square cm for entry 2) Flexible TPS for a deployable aerodynamic decelerator with the capability to withstand dual pulsed heating environments as high as 120 W/square cm for aerocapture and 30 W/square cm for entry NASA, along with its vendors, has begun developing and testing materials for each of the deceleration approaches. These include multi-layer rigid ablators and flexible ablative materials. In order to model the response of these types of materials, new and improved modeling techniques will be required. This presentation will outline the types of materials that are under development and illustrate the need for advancement in modeling of ablative materials.

Beck, Robin A.↗
NASA NTRS2011

Overview of the NASA Entry, Descent and Landing Systems Analysis Exploration Feed-Forward Study

Technology required to land large payloads (20 to 50 mt) on Mars remains elusive. In an effort to identify the most viable investment path, NASA and others have been studying various concepts. One such study, the Entry, Descent and Landing Systems Analysis (EDLSA) Study [1] identified three potential options: the rigid aeroshell, the inflatable aeroshell and supersonic retropropulsion (SRP). In an effort to drive out additional levels of design detail, a smaller demonstrator, or exploration feed-forward (EFF), robotic mission was devised that utilized two of the three (inflatable aeroshell and SRP) high potential technologies in a configuration to demonstrate landing a two to four metric ton payload on Mars. This paper presents and overview of the maximum landed mass, inflatable aeroshell controllability and sensor suite capability assessments of the selected technologies and recommends specific technology areas for additional work.

DwyerCianciolo, Alicia M.↗
NASA NTRS2012

Entry, Descent, and Landing for Human Mars Missions

One of the most challenging aspects of a human mission to Mars is landing safely on the Martian surface. Mars has such low atmospheric density that decelerating large masses (tens of metric tons) requires methods that have not yet been demonstrated, and are not yet planned in future Mars missions. To identify the most promising options for Mars entry, descent, and landing, and to plan development of the needed technologies, NASA's Human Architecture Team (HAT) has refined candidate methods for emplacing needed elements of the human Mars exploration architecture (such as ascent vehicles and habitats) on the Mars surface. This paper explains the detailed, optimized simulations that have been developed to define the mass needed at Mars arrival to accomplish the entry, descent, and landing functions. Based on previous work, technology options for hypersonic deceleration include rigid, mid-L/D (lift-to-drag ratio) aeroshells, and inflatable aerodynamic decelerators (IADs). The hypersonic IADs, or HIADs, are about 20% less massive than the rigid vehicles, but both have their technology development challenges. For the supersonic regime, supersonic retropropulsion (SRP) is an attractive option, since a propulsive stage must be carried for terminal descent and can be ignited at higher speeds. The use of SRP eliminates the need for an additional deceleration system, but SRP is at a low Technology Readiness Level (TRL) in that the interacting plumes are not well-characterized, and their effect on vehicle stability has not been studied, to date. These architecture-level assessments have been used to define the key performance parameters and a technology development strategy for achieving the challenging mission of landing large payloads on Mars.

Munk, Michelle M.↗
NASA NTRS2012

Compilation of Abstracts for SC12 Conference Proceedings

1 A Breakthrough in Rotorcraft Prediction Accuracy Using Detached Eddy Simulation; 2 Adjoint-Based Design for Complex Aerospace Configurations; 3 Simulating Hypersonic Turbulent Combustion for Future Aircraft; 4 From a Roar to a Whisper: Making Modern Aircraft Quieter; 5 Modeling of Extended Formation Flight on High-Performance Computers; 6 Supersonic Retropropulsion for Mars Entry; 7 Validating Water Spray Simulation Models for the SLS Launch Environment; 8 Simulating Moving Valves for Space Launch System Liquid Engines; 9 Innovative Simulations for Modeling the SLS Solid Rocket Booster Ignition; 10 Solid Rocket Booster Ignition Overpressure Simulations for the Space Launch System; 11 CFD Simulations to Support the Next Generation of Launch Pads; 12 Modeling and Simulation Support for NASA's Next-Generation Space Launch System; 13 Simulating Planetary Entry Environments for Space Exploration Vehicles; 14 NASA Center for Climate Simulation Highlights; 15 Ultrascale Climate Data Visualization and Analysis; 16 NASA Climate Simulations and Observations for the IPCC and Beyond; 17 Next-Generation Climate Data Services: MERRA Analytics; 18 Recent Advances in High-Resolution Global Atmospheric Modeling; 19 Causes and Consequences of Turbulence in the Earths Protective Shield; 20 NASA Earth Exchange (NEX): A Collaborative Supercomputing Platform; 21 Powering Deep Space Missions: Thermoelectric Properties of Complex Materials; 22 Meeting NASA's High-End Computing Goals Through Innovation; 23 Continuous Enhancements to the Pleiades Supercomputer for Maximum Uptime; 24 Live Demonstrations of 100-Gbps File Transfers Across LANs and WANs; 25 Untangling the Computing Landscape for Climate Simulations; 26 Simulating Galaxies and the Universe; 27 The Mysterious Origin of Stellar Masses; 28 Hot-Plasma Geysers on the Sun; 29 Turbulent Life of Kepler Stars; 30 Modeling Weather on the Sun; 31 Weather on Mars: The Meteorology of Gale Crater; 32 Enhancing Performance of NASAs High-End Computing Applications; 33 Designing Curiosity's Perfect Landing on Mars; 34 The Search Continues: Kepler's Quest for Habitable Earth-Sized Planets.

HPC↗
NASA NTRSDate not supplied

Parametric Studies of Human Mars Entry, Descent, and Landing Systems

This paper will present a parametric analysis for entry, descent, and landing (EDL) concepts, enabling rapid systems assessment and tradespace exploration. The entry system uses a hypersonic inflatable aerodynamic decelerator (HIAD) technology. The baseline system includes elements for Mars aerocapture (AC) and EDL segments of the mission. The mission concept of operations (ConOps) begins at Mars arrival in a polar inclination. After performing an aerocapture (AC) maneuver into a pre-defined Mars parking orbit, the AC HIAD is jettisoned. The rest of thevehicle stays in the parking orbit for up to one year. The EDL sequence starts with a deorbit burn at the apoapsis of the parking orbit. After hypersonic entry with the EDL HIAD, the entry system uses a supersonic retropropulsion maneuver to slow the vehicle for the descent and landing segments of the mission. The vehicle will maintain a constant velocity of 2.5 m/s for 5 seconds prior to landing.The system includes a Mars Ascent Vehicle (MAV), Mars Descent Module (MDM), and two HIADs. The MDM includes a primary structure, tanks, engines, and radiators. The primary structure is an aluminum-lithium (Al-Li)cruciform design similar to the structural design of the Apollo Lunar Module. The cruciform planform layout results in four outer bays, with adequate volume in the corners between outer bays to package four landing gears. The central bay is reserved for packaging the MAV and the recessed MAV engines. Two of the outer bays accommodate main propellant tanks, with one LOX and one CH4 tank in each bay. The two remaining outer bays each house four rocket engine systems and associated support structure. Each HIAD comprises an inflatable structure, flexible thermal protection system, gas, and gas generators. The HIAD design used in this study is a stacked-toroid concept with pairing loop straps and radial/chevron straps. The baseline system lands a 20-t payload on the Mars surface. It is assumed the vehicle arrives at 6.2 km/s relative velocity at 90° inclination and is captured to a 1-Sol parking orbit.The baseline design includes many assumptions such as margins, arrival state, ConOps options, parking orbit, physical dimensions, propellant options, and technology concepts. The impact of these parameters are quantified through systems-level sensitivity analyses, which capture the global impact—not at a component level—but at the systems level. The systems-level sensitivities expose major design drivers and importance of each assumption for a design.Through tradespace exploration, a wide range of systems parameters are examined and compared for several feasible design options. Studies have been completed for the following input parameters: payload mass, propellant options, AC/EDL ballistic coefficient, lander thrust to weight ratio (T/W) (surrogate for the maximum EDL g’s),engine specific impulse (Isp), parking orbit, and inclination.The final paper will present and discuss the parametric approach used in the study. It will also include the results of recent systems analyses, sensitivity analyses, and tradespace exploration

Jamshid Samareh↗
NASA NTRSDate not supplied

Preparations for Tomographic Background-Oriented Schlieren at the 31-Inch Mach 10 Wind Tunnel

This paper details effort by researchers at NASA Langley Research Center to prepare for an upcoming high-speed tomographic background-oriented schlieren test entry in the 31-Inch Mach 10 wind tunnel. In order to troubleshoot potential problems that will arise when setting up the experiment in the actual test facility, a full-scale mock-up of the 31-Inch Mach 10 wind tunnel test section was constructed in the laboratory to provide equivalent characteristic optical access and mounting options. A 3D-printed model that replicates the scale of the forebody of a supersonic retropropulsion (SRP) model that will be used in the final test entry has been fabricated with a central nozzle that can be connected to a high-pressure gas supply to approximate the nozzle flow that will occur. Six high-speed imaging systems consisting of a high-speed camera with a fiber illuminator module have been assembled and mounted on the test section mock-up to provide six unique perspective views of the nozzle flow from the model. Tests with one of the imaging systems using a pulsed laser light source showed that sufficient signal intensity and illumination uniformity was achieved, as demonstrated by the 2D BOS data obtained with the system. Additional work to show that retroreflective material would survive inside of the test section of the 31-Inch Mach 10 tunnel was also performed. Finally, a remotely-controlled tomographic BOS calibration system is discussed.

Brett F Bathel↗
NASA NTRSDate not supplied

Novel Turbulence Model Tailored for Complex Rocket Motor Exhaust Jets

This paper describes a new RANS turbulence model intended to give improved accuracy in turbulence that is affected by compressibility, rotation and curvature, and variable density, making it well-suited to complex rocket motor exhaust jets. Four key elements of the model are: the use of simple-averaged (not mass-averaged) turbulence decomposition; a second-order Taylor series for the equilibrated turbulent fluxes; empirical modeling that accounts for the combined effects of aforementioned turbulence modifiers; and model tuning on large-eddy simulation data for mixing layers for a wide range of these turbulence modifiers. The model is implemented as an extension of the k-w SST model. Results include comparisons to detached eddy simulation results for hot and cold co-flowing exhaust jets, and to experimental data for supersonic retropropulsion on the CobraMRV Mars-lander concept vehicle. The model is named TMRC for its constituents elements, the Taylor series with Mach number, rotation and curvature effects.

turbulence modeling↗
NASA NTRS2020

Computational Investigation of Powered Descent for Human-Scale Mars Landers

NASA human-scale Mars entry, descent, and landing vehicles currently under consideration have system requirements beyond the current state of the art. Specifically, heritage approaches for deceleration using parachutes and initiating engines at subsonic speeds are not adequate to land these large vehicles (50-60 metric ton entry mass) precisely (within 50 meters of a surface target). Therefore, the present approach for landing humans on Mars utilizes entry vehicles with large aerodynamic surface areas that transition to the descent and landing phase by initiating retropropulsion at supersonic speeds. Vehicle performance, including stability and control, depends on accurate characterization of the aerodynamic-propulsive interference effects, especially at engine initiation, when the vehicles are transitioning from an entry attitude to that suitable for descent and landing. This paper summarizes the human Mars entry vehicle configurations, retropropulsion integration, performance assumptions, and expected flight environment at engine initiation. Cases spanning supersonic, transonic, and subsonic conditions during powered descent are analyzed using four computational fluid dynamics codes, all of which have been previously applied to supersonic retropropulsion at varying scales and conditions. The results of the analysis are used to develop an aerodynamic-propulsive interference model that has since been applied in trajectory simulations to augment the fidelity of human-scale Mars entry vehicle performance.

Ashley M Korzun↗
NASA NTRSDate not supplied

Computational Modeling of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. A test will be conducted in the NASA Langley Unitary Plan Wind Tunnel to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. This paper covers pre-test computational flowfield predictions of two different models derived from full-scale reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. Calculations of the blunt model include variations in nozzle configuration: nozzle location, size, area ratio, and pointing direction. There exist some significant differences between solvers, but some general trends are observed from simulations of the blunt model. First, aerodynamic axial force from the heatshield decreases with increasing thrust due to expanding plume blockage. Second, nozzles that point along the model axis result in lower aerodynamic axial force compared to nozzles that have a radial thrust component. Finally, placing the nozzles further away from the model nose preserves more heatshield axial force with increasing thrust com- pared to nozzles that are closer to the nose. For the slender model, the axial force from the heatshield is similar to the non-blowing axial force regardless of thrust magnitude, due to the nozzle arrangement on the heatshield. Once the test is completed, direct comparisons between the computations and test data will be made to determine computational uncertainties in a wind tunnel environment, to identify gaps in predictive capabilities, and to inform planning for future ground and flight test programs for Mars powered descent vehicles.

Mars↗
NASA NTRSDate not supplied

Computational Modeling of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. A test will be conducted in the NASA Langley Unitary Plan Wind Tunnel to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. This paper covers pre-test computational flowfield predictions of two different models derived from full-scale reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. Calculations of the blunt model include variations in nozzle configuration: nozzle location, size, area ratio, and pointing direction. There exist some significant differences between solvers, but some general trends are observed from simulations of the blunt model. First, aerodynamic axial force from the heatshield decreases with increasing thrust due to expanding plume blockage. Second, nozzles that point along the model axis result in lower aerodynamic axial force compared to nozzles that have a radial thrust component. Finally, placing the nozzles further away from the model nose preserves more heatshield axial force with increasing thrust compared to nozzles that are closer to the nose. For the slender model, the axial force from the heatshield is similar to the non-blowing axial force regardless of thrust magnitude, due to the nozzle arrangement on the heatshield. Once the test is completed, direct comparisons between the computations and test data will be made to determine computational uncertainties in a wind tunnel environment, to identify gaps in predictive capabilities, and to inform planning for future ground and flight test programs for Mars powered descent vehicles.

Supersonic Retropropulsion↗
NASA NTRSDate not supplied

Status of Mars Retropropulsion Testing in the Langley Unitary Plan Wind Tunnel

Future Mars human landings will be enabled by a powered descent phase starting at supersonic conditions, something which has never been done before on a Mars mission. Significant aerosciences challenges exist due to jet interactions between the retrorocket engine plumes, freestream flow, and vehicle that will affect the aerodynamic behavior during powered descent. Historically, wind tunnel tests have been used to study the interactions with inert gas exhaust simulants in place of rocket engines. On the computational side, flowfield simulations have been completed at full-scale conditions, but the available ground and flight data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles, due to insufficient data, dissimilar vehicle geometries, and disparate operating conditions. A wind tunnel test has been designed to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. The test will be conducted in the NASA Langley Unitary Plan Wind Tunnel and is designed with improvements in model design and data products over past tests. The test campaign will be run using sub-scale model geometries derived from NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry that generates higher unpowered lift. Both models have been fabricated and are ready for testing. The blunt model is equipped with the flexibility to examine the effects of nozzle pointing direction, number, location, size, and area ratio. The main measurements are heatshield aerodynamic interference forces and moments with a custom flow-through balance, discrete and distributed heatshield pressure, and high-speed flowfield visualization. This paper covers the test objectives, facility, models and instrumentation, and planned test matrix.

Mars↗
NASA NTRSDate not supplied

Status of Mars Retropropulsion Testing in the Langley Unitary Plan Wind Tunnel

Future Mars human landings will be enabled by a powered descent phase starting at supersonic conditions, something which has never been done before on a Mars mission. Significant aerosciences challenges exist due to jet interactions between the retrorocket engine plumes, freestream flow, and vehicle that will affect the aerodynamic behavior during powered descent. Historically, wind tunnel tests have been used to study the interactions with inert gas exhaust simulants in place of rocket engines. On the computational side, flowfield simulations have been completed at full-scale conditions, but the available ground and flight data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles, due to insufficient data, dissimilar vehicle geometries, and disparate operating conditions. A wind tunnel test has been designed to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. The test will be conducted in the NASA Langley Unitary Plan Wind Tunnel and is designed with improvements in model design and data products over past tests. The test campaign will be run using sub-scale model geometries derived from NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry that generates higher unpowered lift. Both models have been fabricated and are ready for testing. The blunt model is equipped with the flexibility to examine the effects of nozzle pointing direction, number, location, size, and area ratio. The main measurements are heatshield aerodynamic interference forces and moments with a custom flow-through balance, discrete and distributed heatshield pressure, and high-speed flowfield visualization. This paper covers the test objectives, facility, models and instrumentation, and planned test matrix.

Supersonic Retropropulsion↗
NASA NTRS2023

Passive Rocket Diffuser Testing: Performance Similitude Via Nozzle Contour Modification

Subscale rocket testing offers a chance to examine exhaust plume behavior and prove out auxiliary systems early in a development cycle, when the cost of design iteration remains low. Despite the utility of experimentation, perfect similitude between a rocket engine and a scaled model is unobtainable; choices must be made to prioritize the reproduction of certain parameters at the expense of others. The experimentalist’s toolkit contains a variety of techniques to achieve this selective similarity. Determination of the appropriate tool is dependent on a test series’ specific objectives, facility limitations, and financial considerations. Preservation of full nozzle geometry is crucial if its internal flow is of interest. However, subscale thrusters often serve as a simple plume source in the evaluation of adjacent phenomena: vehicle base pressures, retropropulsive flows, supersonic diffuser performance, refractory erosion, deflector cooling, impingement shock structures, etc. In such cases, it may be permissible to rework the nozzle geometry to compensate for discrepancies in subscale propellant pressure, temperature, or chemical composition. Thrust, pressure, Mach number, heat flux, and myriad other parameters can be used to ground the altered designs. Subscale diffuser testing at the full chamber pressure of flight hardware is occasionally infeasible. It is hypothesized that, absent a means of backpressure reduction, comparable performance may be obtained for a prescribed motive fluid by maintaining scaled engine mass flow and nozzle exit geometry at the expense of throat geometry, exit pressure, and Mach number. The present work offers discrete empirical support for this mass-flow-based approach to contour modification.

subscale↗