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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.

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Modeling Powered Aerodynamics for the Orion Launch Abort Vehicle Aerodynamic Database

Modeling the aerodynamics of the Orion Launch Abort Vehicle (LAV) has presented many technical challenges to the developers of the Orion aerodynamic database. During a launch abort event, the aerodynamic environment around the LAV is very complex as multiple solid rocket plumes interact with each other and the vehicle. It is further complicated by vehicle separation events such as between the LAV and the launch vehicle stack or between the launch abort tower and the crew module. The aerodynamic database for the LAV was developed mainly from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamic simulations. However, limitations in both methods have made it difficult to properly capture the aerodynamics of the LAV in experimental and numerical simulations. These limitations have also influenced decisions regarding the modeling and structure of the aerodynamic database for the LAV and led to compromises and creative solutions. Two database modeling approaches are presented in this paper (incremental aerodynamics and total aerodynamics), with examples showing strengths and weaknesses of each approach. In addition, the unique problems presented to the database developers by the large data space required for modeling a launch abort event illustrate the complexities of working with multi-dimensional data.

Chan, David T.↗
NASA NTRS1991

Navier-Stokes computations with finite-rate chemistry for LO2/LH2 rocket engine plume flow studies

Computational fluid dynamics methods have been developed and applied to Space Shuttle Main Engine LO2/LH2 plume flow simulation/analysis of airloading and convective base heating effects on the vehicle at high flight velocities and altitudes. New methods are described which were applied to the simulation of a Return-to-Launch-Site abort where the vehicle would fly briefly at negative angles of attack into its own plume. A simplified two-perfect-gases-mixing approach is used where one gas is the plume and the other is air at 180-deg and 135-deg flight angle of attack. Related research has resulted in real gas multiple-plume interaction methods with finite-rate chemistry described herein which are applied to the same high-altitude-flight conditions of 0 deg angle of attack. Continuing research plans are to study Orbiter wake/plume flows at several Mach numbers and altitudes during ascent and then to merge this model with the Shuttle 'nose-to-tail' aerodynamic and SRB plume models for an overall 'nose-to-plume' capability. These new methods are also applicable to future launch vehicles using clustered-engine LO2/LH2 propulsion.

Dougherty, N. Sam↗
NASA NTRS1993

Aerodynamics of powered missile separation from F/A-18 aircraft

A 3D dynamic 'chimera' algorithm that solves the thin-layer Navier-Stokes equations over multiple moving bodies was modified to numerically simulate the aerodynamics, missile dynamics, and missile plume interactions of a missile separating from a generic wing and from an F/A-18 aircraft in transonic flow. The missile is mounted below the wing for missile separation from the wing and on the F/A-18 fuselage at the engine inlet side for missile separation from aircraft. Static and powered missile separation cases are considered to examine the influence of the missile and plume on the wing and F/A-18 fuselage and engine inlet. The aircraft and missile are at two degrees angle of attack, Reynolds number of 10 million, freestream Mach number of 1.05 and plume Mach number of 3.0. The computational results show the details of the flow field.

Ahmad, J. U.↗
NASA NTRS2007

Reaction Control System Design Considerations for Mars Entry Vehicles

The next generation of Mars exploration landers must precisely deliver scientific payloads to sites of interest, unlike previous Mars missions. The past missions, such as Viking and Pathfinder, performed landings to within 100s of kilometers from their targets using an unguided atmospheric entry. Guided entry of a capsule with a relatively high lift-to-drag ratio will allow landing to within 10s of kilometers from the target with a significantly more massive payload. Successful guided entry requires the use of a reaction control system (RCS) for both attitude correction and entry guidance maneuvers. Various aspects of the entry, descent and landing (EDL) system performance may be impacted by the operation of the RCS during entry. This paper illustrates the risks that arise from the gasdynamic interaction of the entry vehicle (EV) and RCS, and which require attention in the areas of aerodynamics and control, and aerothermal environments. This paper will review the methods to address the design challenges associated with integration of RCS into the atmospheric entry system. Among these challenges is the analysis of the potential for the aerodynamic interference due to both the direct jet plume impingement and more complex plume interactions with the wake flow. These interactions can result in enhanced aeroheating, requiring that a different approach to the thermal protection system (TPS) selection and sizing be used. The recent findings for Mars Science Laboratory and Mars Phoenix will be presented to help illustrate some of the phenomena. Current design solutions will be discussed.

Dyakonov, Artem A.↗
NASA NTRS2012

Overheating Anomalies during Flight Test Due to the Base Bleeding

In this paper we present the results of the analytical and numerical studies of the plume interaction with the base flow in the presence of base out-gassing. The physics-based analysis and CFD modeling of the base heating for single solid rocket motor performed in this research addressed the following questions: what are the key factors making base flow so different from that in the Shuttle [1]; why CFD analysis of this problem reveals small plume recirculation; what major factors influence base temperature; and why overheating was initiated at a given time in the flight. To answer these questions topological analysis of the base flow was performed and Korst theory was used to estimate relative contributions of radiation, plume recirculation, and chemically reactive out-gassing to the base heating. It was shown that base bleeding and small base volume are the key factors contributing to the overheating, while plume recirculation is effectively suppressed by asymmetric configuration of the flow formed earlier in the flight. These findings are further verified using CFD simulations that include multi-species gas environment both in the plume and in the base. Solid particles in the exhaust plume (Al2O3) and char particles in the base bleeding were also included into the simulations and their relative contributions into the base temperature rise were estimated. The results of simulations are in good agreement with the temperature and pressure in the base measured during the test.

base bleeding↗
NASA NTRS1975

Space shuttle orbiter reaction control system jet interaction study

The space shuttle orbiter has forward mounted and rear mounted Reaction Control Systems (RCS) which are used for orbital maneuvering and also provide control during entry and abort maneuvers in the atmosphere. The effects of interaction between the RCS jets and the flow over the vehicle in the atmosphere are studied. Test data obtained in the NASA Langley Research Center 31 inch continuous flow hypersonic tunnel at a nominal Mach number of 10.3 is analyzed. The data were obtained with a 0.01 scale force model with aft mounted RCS nozzles mounted on the sting off of the force model balance. The plume simulations were accomplished primarily using air in a cold gas simulation through scaled nozzles, however, various cold gas mixtures of Helium and Argon were also tested. The effect of number of nozzles was tested as were limited tests of combined controls. The data show that RCS nozzle exit momentum ratio is the primary correlating parameter for effects where the plume impinges on an adjacent surface and mass flow ratio is the parameter where the plume interaction is primarily with the external stream. An analytic model of aft mounted RCS units was developed in which the total reaction control moments are the sum of thrust, impingement, interaction, and cross-coupling terms.

Rausch, J. R.↗
NASA NTRSDate not supplied

Large Vehicle Lunar Landing Surface Interaction and In-Situ Resource Based Risk Mitigation: Landing & Launch Pads

A key capability required for the exploration of planetary bodies is the ability to land on the surface. Previous work performed by NASA and other institutions has primarily focused on landing small spacecraft on planetary surfaces and the associated small-to-medium thrusters required for the soft landing. In the case of human exploration—particularly the establishment of long duration exploration and habitation outposts—the ability to land large landers, such as the SpaceX Starship, is necessary. These larger landing systems require the use of more powerful engines, with higher engine exhaust temperatures and higher landing loads. Understanding the excavation of material by the engines, as well as the potential for the landing legs to sink into the subsurface, is key in ensuring reliable and safe landings. A further improvement in landing reliability can be achieved by constructing landing / launch pads, especially with in-situ resources. Some material excavation by the plume is inevitable, leaving at least a portion of the surface scoured and uneven under the lander and ejecting regolith particles and rocks at very high velocities. One possible solution would be to robotically build landing / launch pads (ideally autonomously) at the destination using in-situ materials. In this case, the first one or few landers will need to land on unimproved surfaces at higher risk; however, they would bring the required equipment to build the landing pads with mostly local resources, thus increasing the reliability of safe landing for subsequent larger landers. A number of methods to build in-situ landing and launch pads have already been developed. These methods include no, or some, addition of required binder additives to the local regolith material, different processing approaches and result in varying landing pad strengths. A sub-scale rocket engine plume, was used to simulate some of the conditions of a landing on the Moon to assess the effectiveness of various materials for an in-situ built landing pad, The GO2/GCH4rocket engine fired on a 1m2area coupons of representative pad materials. The results will allow continued development towards materials that satisfy the landing pad properties required for the effective risk reduction and increased reliability for landing people and equipment on the lunar surface. This work contained two parts: (1) computer modeling of a large rocket engine plume interacting with regolith on the Moon, using the Granular Gas Flow Solver (GGFS) provided by CFD Research Corporation as well as other computational fluid dynamics codes (CFD) such as Loci/CHEM. (2) Developing landing/launch pad materials that could be used for in-situ construction on the lunar surface in the future, to mitigate the calculated effects of a large vehicle rocket engine landing and launching on the Moon.

Lunar↗
NASA NTRS1993

Four-Nozzle Benchmark Wind Tunnel Model USA Code Solutions for Simulation of Multiple Rocket Base Flow Recirculation at 145,000 Feet Altitude

Multiple rocket exhaust plume interactions at high altitudes can produce base flow recirculation with attendant alteration of the base pressure coefficient and increased base heating. A search for a good wind tunnel benchmark problem to check grid clustering technique and turbulence modeling turned up the experiment done at AEDC in 1961 by Goethert and Matz on a 4.25-in. diameter domed missile base model with four rocket nozzles. This wind tunnel model with varied external bleed air flow for the base flow wake produced measured p/p(sub ref) at the center of the base as high as 3.3 due to plume flow recirculation back onto the base. At that time in 1961, relatively inexpensive experimentation with air at gamma = 1.4 and nozzle A(sub e)/A of 10.6 and theta(sub n) = 7.55 deg with P(sub c) = 155 psia simulated a LO2/LH2 rocket exhaust plume with gamma = 1.20, A(sub e)/A of 78 and P(sub c) about 1,000 psia. An array of base pressure taps on the aft dome gave a clear measurement of the plume recirculation effects at p(infinity) = 4.76 psfa corresponding to 145,000 ft altitude. Our CFD computations of the flow field with direct comparison of computed-versus-measured base pressure distribution (across the dome) provide detailed information on velocities and particle traces as well eddy viscosity in the base and nozzle region. The solution was obtained using a six-zone mesh with 284,000 grid points for one quadrant taking advantage of symmetry. Results are compared using a zero-equation algebraic and a one-equation pointwise R(sub t) turbulence model (work in progress). Good agreement with the experimental pressure data was obtained with both; and this benchmark showed the importance of: (1) proper grid clustering and (2) proper choice of turbulence modeling for rocket plume problems/recirculation at high altitude.

Dougherty, N. S.↗
NASA NTRS1977

Space shuttle orbiter rear mounted reaction control system jet interaction study

The effect of interaction between the reaction control system (RCS) jets and the flow over the space shuttle orbiter in the atmosphere was investigated in the NASA Langley 31-inch continuous flow hypersonic tunnel at a nominal Mach number of 10.3 and in the AEDC continuous flow hypersonic tunnel B at a nominal Mach number of 6, using 0.01 and .0125 scale force models with aft RCS nozzles mounted both on the model and on the sting of the force model balance. The data show that RCS nozzle exit momentum ratio is the primary correlating parameter for effects where the plume impinges on an adjacent surface and mass flow ratio is the parameter when the plume interaction is primarily with the external stream. An analytic model of aft mounted RCS units was developed in which the total reaction control moments are the sum of thrust, impingement, interaction, and cross-coupling terms.

Rausch, J. R.↗
NASA NTRS2016

Modeling an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT)

An iodine-operated 200-W Hall thruster plume has been simulated using a hybrid-PIC model to predict the spacecraft surface-plume interaction for spacecraft integration purposes. For validation of the model, the plasma potential, electron temperature, ion current flux, and ion number density of xenon propellant were compared with available measurement data at the nominal operating condition. To simulate iodine plasma, various collision cross sections were found and used in the model. While time-varying atomic iodine species (i.e., I, I+, I2+) information is provided by HPHall simulation at the discharge channel exit, the molecular iodine species (i.e., I2, I2+) are introduced as Maxwellian particles at the channel exit. Simulation results show that xenon and iodine plasma plumes appear to be very similar under the assumptions of the model. Assuming a sticking coefficient of unity, iodine deposition rate is estimated.

Numerical Simulation↗
NASA NTRS2016

Modeling an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT)

An iodine-operated 200-W Hall thruster plume has been simulated using a hybrid-PIC model to predict the spacecraft surface-plume interaction for spacecraft integration purposes. For validation of the model, the plasma potential, electron temperature, ion current flux, and ion number density of xenon propellant were compared with available measurement data at the nominal operating condition. To simulate iodine plasma, various collision cross sections were found and used in the model. While time-varying atomic iodine species (i.e., I, I+, I2+) information is provided by HP Hall simulation at the discharge channel exit, the molecular iodine species (i.e., I2, I2+) are introduced as Maxwellian particles at the channel exit. Simulation results show that xenon and iodine plasma plumes appear to be very similar under the assumptions of the model. Assuming a sticking coefficient of unity, iodine deposition rate is estimated.

Iodine↗
NASA NTRS2008

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively. The Mach Number distribution shows small yet distinct differences between the two cases such as locations of shocks/shock reflections and a slightly different impingement point on the wall of the diffuser from the expansion at the exit of the nozzle. Similarly the temperature distribution indicates different flow recirculation patterns in the test cell. Both cases capture all the essential flow phenomena such as the shock-boundary layer interaction, plume expansion, expansion of the first stage ejectors, mixing between the engine plume and the first stage ejector flow and pressurization due to the first stage ejectors. The final paper will discuss thermal loads on the walls of the diffuser and cooling mechanisms investigated.

Allgood, Daniel C.↗
NASA NTRS1992

Three dimensional particle simulation of high altitude rocket plumes

The interaction of two nozzles exhausting into vacuum generates a complex three-dimensional shock structure. The shock structure and resulting plume flowfield is characterized by the nozzle separation distance. For the appropriate range of penetration Knudsen numbers, the analysis of this shock structure can be suitably accomplished through a Monte Carlo simulation. This paper describes the application of a general three-dimensional Monte Carlo simulation on the Connection Machine CM-2 to the analysis of the plume self-interaction shock in the near field. Results are presented for two cases, corresponding to a small and a large nozzle separation distance. The results correctly reproduce the expected flow features and demonstrate the ability of this method to properly simulate the start of the plume self-interaction shock. This has significance not only for allowing analysis of the self-interacting plume in the near field, but also for allowing the subsequent simulation of the far field flow through the use of a continuation downstream exit boundary.

Dagum, Leonardo↗
NASA NTRS1976

Investigations of the 0.020-scale 88-OTS Integrated Space Shuttle Vehicle Jet-Plume Model in the NASA/Ames Research Center 11 by11-Foot Unitary Plan Wind Tunnel (IA80)

The results are documented of jet plume effects wind tunnel test of the 0.020-scale 88-OTS launch configuration space shuttle vehicle model in the 11 x 11 foot leg of the NASA/Ames Research Center Unitary Plan Wind Tunnel. This test involved cold gas main propulsion system (MPS) and solid rocket motor (SRB) plume simulations at Mach numbers from 0.6 to 1.4. Integrated vehicle surface pressure distributions, elevon and rudder hinge moments, and wing and vertical tail root bending and torsional moments due to MPS and SRB plume interactions were determined. Nozzle power conditions were controlled per pretest nozzle calibrations. Model angle of attack was varied from -4 deg to +4 deg; model angle of sideslip was varied from -4 deg to +4 deg. Reynolds number was varied for certain test conditions and configurations, with the nominal freestream total pressure being 14.69 psia. Plotted force and pressure data are presented.

Nichols, M. E.↗
NASA NTRS1980

Consolidation of data base for Army generalized missile model

Data from plume interaction tests, nose mounted canard configuration tests, and high angle of attack tests on the Army Generalized Missile model are consolidated in a computer program which makes them readily accessible for plotting, listing, and evaluation. The program is written in FORTRAN and will run on an ordinary minicomputer. It has the capability of retrieving any coefficient from the existing DATAMAN tapes and displaying it in tabular or plotted form. Comparisons of data taken in several wind tunnels and of data with the predictions of Program MISSILE2 are also presented.

Klenke, D. J.↗
NASA NTRS1981

Photochemistry in planetary atmospheres

Widely varying paths of evolutionary history, atmospheric processes, solar fluxes, and temperatures have produced vastly different planetary atmospheres. The similarities and differences between the earth atmosphere and those of the terrestrial planets (Venus and Mars) and of the Jovian planets are discussed in detail; consideration is also given to the photochemistry of Saturn, Uranus, Pluto, Neptune, Titan, and Triton. Changes in the earth's ancient atmosphere are described, and problems of interest in the earth's present troposphere are discussed, including the down wind effect, plume interactions, aerosol nucleation and growth, acid rain, and the fate of terpenes. Temperature fluctuations in the four principal layers of the earth's atmosphere, predicted decreases in the ozone concentration as a function of time, and spectra of particles in the earth's upper atmosphere are also presented. Finally, the vertical structure of the Venus cloud system and the thermal structure of the Jovian planets are shown graphically.

Levine, J. S.↗
NASA NTRS1985

Results of the SPAS-01 RCS plume impingement test

The Orbiter RCS flowfield is modeled in the free molecular region and an impingement analysis is performed to characterize on-orbit plume interactions with nearby bodies, such as the SPAS-01 satellite. The flow in the far field is treated with a source flow model at any point 40 ft from the nozzle. Impingement is calculated according to Newtonian impact theory. The flow density in the far field is expressed in terms of a conservation of energy equation, with account taken of the expansion of the gas, cooling, and the angular distribution. The SPAS-01 platform gathered data at eight different positions around the Orbiter to test the model predictions. Excellent agreement was obtained in predictions of the torque and rotation rate induced on the SPAS by firing the thrusters.

Lazaron, M. P.↗
NASA NTRS1995

A vehicle health monitoring system for the Space Shuttle Reaction Control System during reentry

A general two tier framework for vehicle health monitoring of Guidance Navigation and Control (GN&C) system actuators, effectors, and propulsion devices is presented. In this context, a top level monitor that estimates jet thrust is designed for the Space Shuttle Reaction Control System (RCS) during the reentry phase of flight. Issues of importance for the use of estimation technologies in vehicle health monitoring are investigated and quantified for the Shuttle RCS demonstration application. These issues include rate of convergence, robustness to unmodeled dynamics, sensor quality, sensor data rates, and information recording objectives. Closed loop simulations indicate that a Kalman filter design is sensitive to modeling error and robust estimators may reduce this sensitivity. Jet plume interaction with the aerodynamic flowfield is shown to be a significant effect adversely impacting the ability to accurately estimate thrust.

Rosello, Anthony David↗