An event recorder for obtaining missile heat- shield ablation data.
Event recorder for in-flight measurements of missile heat-shield ablation
HEAT SHIELD↗
SEARCH · Engineering Papers
Results for “Aerodynamic Heating”
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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The effect of real-air properties upon aerodynamic forces, moments, and heat transfer rates for reentry vehicles
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Cohen, N. B.↗
Aerodynamic pressure and heating-rate distributions in tile gaps around chine regions with pressure gradients at a Mach number of 6.6
Surface and gap pressures and heating-rate distributions were obtained for simulated Thermal Protection System (TPS) tile arrays on the curved surface test apparatus of the Langley 8-Foot High Temperature Tunnel at Mach 6.6. The results indicated that the chine gap pressures varied inversely with gap width because larger gap widths allowed greater venting from the gap to the lower model side pressures. Lower gap pressures caused greater flow ingress from the surface and increased gap heating. Generally, gap heating was greater in the longitudinal gaps than in the circumferential gaps. Gap heating decreased with increasing gap depth. Circumferential gap heating at the mid-depth was generally less than about 10 percent of the external surface value. Gap heating was most severe at local T-gap junctions and tile-to-tile forward-facing steps that caused the greatest heating from flow impingement. The use of flow stoppers at discrete locations reduced heating from flow impingement. The use of flow stoppers at discrete locations reduced heating in most gaps but increased heating in others. Limited use of flow stoppers or gap filler in longitudinal gaps could reduce gap heating in open circumferential gaps in regions of high surface pressure gradients.
Hunt, L. Roane↗
Determination of Transient Skin Temperature of Conical Bodies During Short-time, High-speed Flight
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HEAT TRANSMISSION - BODIES↗
Nonlinear theory of a hot-wire anemometer
A theoretical analysis is presented for the hot-wire anemometer to determine the differences in resistance characteristics as given by King's equation for an infinite wire length and those given by the additional considerations of (a) a finite length of wire with heat loss through its ends and (b) heat loss due to a nonlinear function of the temperature difference between the wire and the air.
HEAT TRANSMISSION - GASES & SOLIDS↗
Aerodynamic, unsteady, kinetic and heat loss effects on the dynamics and structure of weakly-burning flames
The first objective of the program is to introduce the meritorious counterflow methodology in microgravity in order to quantify the steady and unsteady characteristics of weakly-burning premixed and diffusion flames for a wide variety of conditions including elevated pressures. Subsequently, through detailed modeling and comparisons with the experimental data, to provide physical insight into the elementary mechanisms controlling the flame response. The configuration offers good control over the parameters of interest and can be modelled closely. The knowledge which will be gained from the counterflow flames will be subsequently used to analyze near-limit phenomena related to other configurations by conducting detailed numerical simulations including multidimensional ones. Among the problems to be analyzed are the downward and upward propagation of near-limit flames in tubes and phenomena observed in spherical and cylindrical geometries.
Egolfopoulos, Fokion N.↗
Summary of the 6th AIAA Propulsion Aerodynamics Workshop Nozzle Test Case: Heated Supersonic Axisymmetric Jets
This paper summarizes findings from the sixth AIAA Propulsion Aerodynamics Workshop (PAW) nozzle test case. The focus of the workshop participants was turbulent supersonic round jets with variation in jet Mach number and jet heating. Three nozzles were utilized with on-design jet Mach numbers of 1.36, 1.63, and 2.00. A heated jet for each of these three nozzles at on-design conditions, a temperature matched on-design condition at Mach 1.63, and an off-design heated condition for the Mach 1.63 nozzle were the five cases examined. Workshop participants utilized Reynolds-averaged Navier-Stokes (RANS) and Large Eddy Simulation (LES) based techniques. Computational fluid dynamics (CFD) solutions were compared with non-intrusive measurements of velocities and temperatures, including both mean values and turbulence statistics. In general, the RANS approaches were unable to capture the trend of reduced mixing as the jet Mach number increased. Incorporation of compressibility corrections improved the agreement with experimental measurements in terms of trends with varying jet Mach number, but also increased the jet potential core lengths to be in worse agreement with data. RANS methods were able to reasonably capture trends of jet heating. The LES-based approaches as a whole were able to capture the effects of compressibility more accurately. Effects of jet heating were captured well by LES. None of the LES approaches introduced disturbances from the wall boundary layers of the nozzle into the jet mixing layer, which would affect the jet flow physics modeling in the initial part of the jet shear layer, but would be very computationally demanding.
supersonic jet↗
Turbine Blade and Endwall Heat Transfer Measured in NASA Glenn's Transonic Turbine Blade Cascade
Higher operating temperatures increase the efficiency of aircraft gas turbine engines, but can also degrade internal components. High-pressure turbine blades just downstream of the combustor are particularly susceptible to overheating. Computational fluid dynamics (CFD) computer programs can predict the flow around the blades so that potential hot spots can be identified and appropriate cooling schemes can be designed. Various blade and cooling schemes can be examined computationally before any hardware is built, thus saving time and effort. Often though, the accuracy of these programs has been found to be inadequate for predicting heat transfer. Code and model developers need highly detailed aerodynamic and heat transfer data to validate and improve their analyses. The Transonic Turbine Blade Cascade was built at the NASA Glenn Research Center at Lewis Field to help satisfy the need for this type of data.
Giel, Paul W.↗
Thermal Analysis and Correlation of the Mars Odyssey Spacecraft's Solar Array During Aerobraking Operations
The Mars Odyssey spacecraft made use of multipass aerobraking to gradually reduce its orbit period from a highly elliptical insertion orbit to its final science orbit. Aerobraking operations provided an opportunity to apply advanced thermal analysis techniques to predict the temperature of the spacecraft's solar array for each drag pass. Odyssey telemetry data was used to correlate the thermal model. The thermal analysis was tightly coupled to the flight mechanics, aerodynamics, and atmospheric modeling efforts being performed during operations. Specifically, the thermal analysis predictions required a calculation of the spacecraft's velocity relative to the atmosphere, a prediction of the atmospheric density, and a prediction of the heat transfer coefficients due to aerodynamic heating. Temperature correlations were performed by comparing predicted temperatures of the thermocouples to the actual thermocouple readings from the spacecraft. Time histories of the spacecraft relative velocity, atmospheric density, and heat transfer coefficients, calculated using flight accelerometer and quaternion data, were used to calculate the aerodynamic heating. During aerobraking operations, the correlations were used to continually update the thermal model, thus increasing confidence in the predictions. This paper describes the thermal analysis that was performed and presents the correlations to the flight data.
Dec, John A.↗
HEAT TRANSFER TO BLUNT AXIALLY SYMMETRIC BODIES
Comparison of heat transfer rates to blunt bodies of revolution & evaluation of effects of body-nose geometry on heat transfer rates
BLUNT BODY↗
Development of Multiobjective Optimization Techniques for Sonic Boom Minimization
A discrete, semi-analytical sensitivity analysis procedure has been developed for calculating aerodynamic design sensitivities. The sensitivities of the flow variables and the grid coordinates are numerically calculated using direct differentiation of the respective discretized governing equations. The sensitivity analysis techniques are adapted within a parabolized Navier Stokes equations solver. Aerodynamic design sensitivities for high speed wing-body configurations are calculated using the semi-analytical sensitivity analysis procedures. Representative results obtained compare well with those obtained using the finite difference approach and establish the computational efficiency and accuracy of the semi-analytical procedures. Multidisciplinary design optimization procedures have been developed for aerospace applications namely, gas turbine blades and high speed wing-body configurations. In complex applications, the coupled optimization problems are decomposed into sublevels using multilevel decomposition techniques. In cases with multiple objective functions, formal multiobjective formulation such as the Kreisselmeier-Steinhauser function approach and the modified global criteria approach have been used. Nonlinear programming techniques for continuous design variables and a hybrid optimization technique, based on a simulated annealing algorithm, for discrete design variables have been used for solving the optimization problems. The optimization procedure for gas turbine blades improves the aerodynamic and heat transfer characteristics of the blades. The two-dimensional, blade-to-blade aerodynamic analysis is performed using a panel code. The blade heat transfer analysis is performed using an in-house developed finite element procedure. The optimization procedure yields blade shapes with significantly improved velocity and temperature distributions. The multidisciplinary design optimization procedures for high speed wing-body configurations simultaneously improve the aerodynamic, the sonic boom and the structural characteristics of the aircraft. The flow solution is obtained using a comprehensive parabolized Navier Stokes solver. Sonic boom analysis is performed using an extrapolation procedure. The aircraft wing load carrying member is modeled as either an isotropic or a composite box beam. The isotropic box beam is analyzed using thin wall theory. The composite box beam is analyzed using a finite element procedure. The developed optimization procedures yield significant improvements in all the performance criteria and provide interesting design trade-offs. The semi-analytical sensitivity analysis techniques offer significant computational savings and allow the use of comprehensive analysis procedures within design optimization studies.
Chattopadhyay, Aditi↗
Spacecraft heat-protection requirements for mars aerodynamic braking.
Spacecraft deceleration upon entry into Mars orbit by aerodynamic braking, considering vehicle shape and velocity effects on entry heating
PLANETARY ENTRY↗
Nasa schedules project fire launch in may
Launching of Project FIRE spacecraft for reentry heating experiments
REENTRY EFFECT↗
Notes on Earth Atmospheric Entry for Mars Sample Return Missions
The entry of sample return vehicles (SRVs) into the Earth's atmosphere is the subject of this document. The Earth entry environment for vehicles, or capsules, returning from the planet Mars is discussed along with the subjects of dynamics, aerodynamics, and heat transfer. The material presented is intended for engineers and scientists who do not have strong backgrounds in aerodynamics, aerothermodynamics and flight mechanics. The document is not intended to be comprehensive and some important topics are omitted. The topics considered in this document include basic principles of physics (fluid mechanics, dynamics and heat transfer), chemistry and engineering mechanics. These subjects include: a) fluid mechanics (aerodynamics, aerothermodynamics, compressible fluids, shock waves, boundary layers, and flow regimes from subsonic to hypervelocity; b) the Earth s atmosphere and gravity; c) thermal protection system design considerations; d) heat and mass transfer (convection, radiation, and ablation); e) flight mechanics (basic rigid body dynamics and stability); and f) flight- and ground-test requirements; and g) trajectory and flow simulation methods.
Rivell, Thomas↗
Summary of the 5th Propulsion Aerodynamics Workshop Nozzle Test Case: Heated Nozzle Exhaust Passing Over A Film-Cooled Plate
This paper summarizes findings from the fifth AIAA Propulsion Aerodynamics Workshop nozzle test case. The experimental configuration examined by workshop participants was a subsonic nozzle with a square exit blowing a heated exhaust over a film-cooled plate. Computational fluid dynamics solutions were obtained and compared to experimental measurements of flow velocities and temperatures, as well as plate surface temperatures. The heated nozzle operated with a Mach 0.3 exit flow at a static temperature ratio of 2.7. Cooling air blowing ratios of 0, 1, and 2 were considered. Computational meshes for the flow domain were provided for participants. Workshop participants from eight separate organizations represented government, industry, and academia. A variety of flow solutions were obtained: Most of the flow solutions employed a Reynolds-averaged Navier-Stokes (RANS) approach, but scale-resolving simulations were used in some cases, including wall-modeled large-eddy simulation (LES), and a few hybrid RANS-LES approaches. One Lattice-Boltzmann solver was employed. For heat transfer to the test article, many participants used a conjugate heat transfer approach. Effects of mesh sensitivity, flow solution approach, and wall heat transfer approach are considered. In general, a fully three-dimensional conjugate heat transfer approach enabled better prediction of surface temperatures than simpler wall temperature boundary treatments. Also, the scale resolving approaches provided better prediction of static temperatures in the flow immediately above the plate consisting of the hot jet exhaust boundary layer interacting with the cooling film. Comparisons of plate surface friction drag and heat transfer obtained from the computations are also presented.
nozzle↗
Summary of the 5th Propulsion Aerodynamics Workshop Nozzle Test Case: Heated Nozzle Exhaust Passing Over A Film-Cooled Plate
This paper summarizes findings from the fifth AIAA Propulsion Aerodynamics Workshop nozzle test case. The experimental configuration examined by workshop participants was a subsonic nozzle with a square exit blowing a heated exhaust over a film-cooled plate. Computational fluid dynamics solutions were obtained and compared to experimental measurements of flow velocities and temperatures, as well as plate surface temperatures. The heated nozzle operated with a Mach 0.3 exit flow at a static temperature ratio of 2.7. Cooling air blowing ratios of 0, 1, and 2 were considered. Computational meshes for the flow domain were provided for participants. Workshop participants from eight separate organizations represented government, industry, and academia. A variety of flow solutions were obtained: Most of the flow solutions employed a Reynolds-averaged Navier-Stokes (RANS) approach, but scale-resolving simulations were used in some cases, including wall-modeled large-eddy simulation (LES), and a few hybrid RANS-LES approaches. One Lattice-Boltzmann solver was employed. For heat transfer to the test article, many participants used a conjugate heat transfer approach. Effects of mesh sensitivity, flow solution approach, and wall heat transfer approach are considered. In general, a fully three-dimensional conjugate heat transfer approach enabled better prediction of surface temperatures than simpler wall temperature boundary treatments. Also, the scale resolving approaches provided better prediction of static temperatures in the flow immediately above the plate consisting of the hot jet exhaust boundary layer interacting with the cooling film. Comparisons of plate surface friction drag and heat transfer obtained from the computations are also presented.
nozzle↗
Navier-Stokes analysis of turbine flowfield and external heat transfer
An explicit 2D Navier-Stokes code has been modified and used to analyze the aerodynamics and heat transfer of a transonic turbine cascade. This code is based on a four-stage Runge-Kutta scheme. An algebraic Reynolds stress model (ARSM) and two versions of low Reynolds number (LRN) two-equation turbulence models, Chien's (1982) LRN k-epsilon model and Coakley's (1983) LRN q-omega model, have been employed in the computations. The surface pressure distributions and wake profiles are predicted well by all the models. The k-epsilon model and the k-epsilon/ARSM model yield better predictions of heat transfer than the q-omega model. The k-epsilon/ARSM solutions show some significant, though not dramatic, differences in the predicted skin friction coefficients, heat transfer rates, and performance parameters, as compared to the k-epsilon model. The predicted semiwake width is consistent with the measurement and correlation.
Luo, J.↗
On the possibility of precipitation-induced vertical lightning in the protoplanetary nebula
Many primitive meteorites contain high concentrations of chondrules - millimeter-sized beads of glassy silicate - in a matrix of essentially unprocessed material. These chondrules were apparently melted by some transient high-energy event, then cooled rapidly on timescales of minutes to hours. Among the phenomena that have been suggested as driving mechanisms for the heating event are: (1) high-velocity impacts; (2) magnetic flares; (3) aerodynamic drag heating; and (4) lightning. Although nebular lightning has been made of the efficiency of such lightning in melting chondrule precursors, a quantitative treatment of processes that might produce nebular lightning has not been attempted. We have developed a one-dimensional nebular lightning model that utilizes a spectrum of particle sizes. This model has also been applied to the Earth as a test of its predictive ability. Particles grow to an initial size distribution, then undergo elastic charge-separating collisions. Charge separation is opposed by the mobile free charge in the nebula, which can either accumulate on and neutralize the charged ice particles or move to cancel the large scale electric field. Since the free charge density depends on the grain size distribution, it is determined self-consistently as the particle distribution changes.
Gibbard, S. G.↗