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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Leading-edge receptivity for blunt-nose bodies

This research program investigates boundary-layer receptivity in the leading-edge region for bodies with blunt leading edges. Receptivity theory provides the link between the unsteady distrubance environment in the free stream and the initial amplitudes of the instability waves in the boundary layer. This is a critical problem which must be addressed in order to develop more accurate prediction methods for boundary-layer transition. The first phase of this project examines the effects of leading-edge bluntness and aerodynamic loading for low Mach number flows. In the second phase of the project, the investigation is extended to supersonic Mach numbers. Singular perturbation techniques are utilized to develop an asymptotic theory for high Reynolds numbers. In the first year, the asymptotic theory was developed for leading-edge receptivity in low Mach number flows. The case of a parabolic nose is considered. Substantial progress was made on the Navier-Sotkes computations. Analytical solutions for the steady and unsteady potential flow fields were incorporated into the code, greatly expanding the types of free-stream disturbances that can be considered while also significantly reducing the the computational requirements. The time-stepping algorithm was modified so that the potential flow perturbations induced by the unsteady pressure field are directly introduced throughout the computational domain, avoiding an artificial 'numerical diffusion' of these from the outer boundary. In addition, the start-up process was modified by introducing the transient Stokes wave solution into the downstream boundary conditions.

Kerschen, Edward J.↗
NASA NTRS1992

Leading-edge receptivity for blunt-nose bodies

Boundary-layer receptivity in the leading edge region for bodies with blunt leading edges is investigated in this research program. Receptivity theory provides the link between the unsteady disturbance environment in the freestream and the initial amplitudes of instability waves in the boundary layer. This is a critical problem which must be addressed in order to develop more accurate prediction methods for boundary-layer transition.

Hammerton, P. W.↗
NASA NTRS1978

Numerical solution of the Navier-Stokes equations for blunt nosed bodies in supersonic flows

A time dependent, two dimensional Navier-Stokes code employing the method of body fitted coordinate technique was developed for supersonic flows past blunt bodies of arbitrary shapes. The bow shock ahead of the body is obtained as part of the solution, viz., by shock capturing. A first attempt at mesh refinement in the shock region was made by using the forcing function in the coordinate generating equations as a linear function of the density gradients. The technique displaces a few lines from the neighboring region into the shock region. Numerical calculations for Mach numbers 2 and 4.6 and Reynolds numbers from 320 to 10,000 were performed for a circular cylinder with and without a fairing. Results of Mach number 4.6 and Reynolds number 10,000 for an isothermal wall temperature of 556 K are presented in detail.

Warsi, Z. U. A.↗
NASA NTRSDate not supplied

Recent Langley Aerothermodynamic Laboratory Testing (2019-2022)

In support of flight programs, multiple wind tunnel entries were conducted on reentry vehicle configurations to investigate blunt body aeroheating environments. Blunt body block thermal protection system (TPS) roughness, the Low-Earth Orbit Flight Test of an Inflatable Decelerator aeroshell scalloping, and a woven TPS all were tested to determine the effects of those parameters on increased surface heating to the vehicle. Commercial Crew Program studies also utilized the speed and ease of testing in the LAL to analyze parametric databases of low-earth reentry configurations, like the Boeing CST-100, and vehicles designed to travel to the moon or Mars, such as the SpaceX Starship. LAL wind tunnel investigations also supported several DoD and AFRL flight projects, such as HIFiRE 6 and BOLT. External customers leveraged the LAL testing capability to advance their research programs, including UPSTAR, RocketLab, and Lawrence Livermore National Laboratory. Academia groups benefitted from the inexpensive cost to test in the tunnels for many university collaborations, such as studies conducted in partnership with the University of Maryland on a flat plate, using a pressure rake and investigating fluid structure interactions. Multiple studies in collaboration with Purdue considered boundary layer transition on various cone configurations. Another study was conducted using a cone with shaped roughness elements borrowed from the Arnold Engineering Development Complex. Internal testing included studies using a phosphor thermography system that can measure higher temperatures than the current LAL phosphor capability and a characterization study of the noise in the wind tunnel flow. The breadth of testing across research and commercial paradigms, both internal and external to NASA, different configurations, a variety of test techniques and aerodynamic and aerothermal environments will be discussed in this overview presentation of the recent work conducted in the LAL.

wind tunnel test↗
NASA NTRS1980

Complete viscous flowfield solutions about a blunt parabolic body in a supersonic stream

A hybrid computational technique which splits the flowfield into inviscid and viscous regions is used to investigate the complete flowfield about axisymmetric parabolic blunt bodies in a supersonic stream. The solutions are carried out on the CDC CYBER-203 computer which, with its extensive memory, allows for the use of a large number of finite-difference mesh points, allowing resolution of important flowfield features. A range of freestream Mach number of 2-5 and a range of Re number based on nose radius of 500-125,000 was run for a sonic corner body. Contour plots of density, pressure, and Mach number, velocity vector plots, and surface distributions of pressure, heat transfer, and shear stress are presented. Also, correlations of the downstream extent of the base recirculation region with Re number based on nose radius are given.

Weilmuenster, K. J.↗
NASA NTRS1975

Computation of nonequilibrium three-dimensional inviscid flow over blunt-nosed bodies flying at supersonic speeds

A computer code based on the method of characteristics is described and applied to the study of two- and three-dimensional chemical nonequilibrium flow over sharp and blunt-nosed bodies. Nonequilibrium flow over a wedge is used to show the approach to equilibrium flow, and to demonstrate the nature of the reaction zone behind the bow shock wave. The structure and development of a blunt-body entropy layer in nonequilibrium flow is examined for a blunt cone at zero incidence. Three-dimensional computations for the space shuttle body at 30 deg angle of attack are presented. A nondimensional scaling parameter, the Damkoehler number, which is the ratio of flow time to chemical reaction time, is calculated and its significance discussed.

Rakich, J. V.↗
NASA NTRS1979

Aerodynamic heating on complex configurations

A relatively simple method was developed previously (DeJarnette and Hamilton, 1973; 1975) to calculate laminar transitional and turbulent heating rates on blunt bodies at angles of attack, taking into account entropy-swelling effects. The method has now been extended to include complex configurations and the use of input surface pressure data when available. Applications of the extended method to the HL-10 Lifting Body, a blunted tangent ogive, a blunted tangent-ogive cylinder, and the Viking Aeroshell are discussed.

Dejarnette, F. R.↗
NASA NTRS1980

A vectorized code for calculating laminar and turbulent hypersonic flows about blunt axisymmetric bodies at zero and small angles of attack

A user's guide is provided for a computer code which calculates the laminar and turbulent hypersonic flows about blunt axisymmetric bodies, such as spherically blunted cones, hyperboloids, etc., at zero and small angles of attack. The code is written in STAR FORTRAN language for the CDC-STAR-100 computer. Time-dependent, viscous-shock-layer-type equations are used to describe the flow field. These equations are solved by an explicit, two-step, time asymptotic, finite-difference method. For the turbulent flow, a two-layer, eddy-viscosity model is used. The code provides complete flow-field properties including shock location, surface pressure distribution, surface heating rates, and skin-friction coefficients. This report contains descriptions of the input and output, the listing of the program, and a sample flow-field solution.

Kumar, A.↗
NASA NTRS2019

Full Facility Shock Frame Simulations of the Electric Arc Shock Tube

Radiative heating computations are performed for high speed lunar return experiments conducted in the Electric Arc Shock Tube (EAST) facility at NASA Ames Research Center. The nonequilibrium radiative transport equations are solved via NASA's in-house radiation code NEQAIR using flow field input from US3D flow solver. The post-shock flow properties for the 10 km/s Earth entry conditions are computed using the stagnation line of a blunt-body and a full facility CFD (Computational Fluid Dynamics) simulation of the EAST shock tube. The shocked gas in the blunt-body flow achieves a thermochemical equilibrium away from the shock front whereas EAST flow exhibits a nonequilibrium behavior due to strong viscous dissipation of the shock by boundary layer. The full-tube flow calculations capture the influence of the boundary layer on the shocked gas state and provide a realistic fluid dynamic input for the radiative predictions. The integrated radiance behind the shock is calculated in NEQAIR for wavelength regimes from Vacuum-UltraViolet (VUV) to InfraRed (IR), which are pertinent to the emission characteristics of high enthalpy shock waves in air. These radiance profiles are validated against corresponding EAST shots. The full-tube simulations successfully predict a sharp radiance peak at the shock front which gets smeared in the test data due to the spatial resolution in the measurements. The full facility based radiance behind the shock shows a slightly better match with the test data in the VUV and Red spectral regions, as compared to that from a blunt-body based predictions. The UV radiance is very similar for both geometries and under-predicts the test behavior. The IR test data matches better with the blunt-body based predictions where the full-tube simulations show a significant over-prediction.

Radiation↗
NASA NTRS1974

Numerical solution of the steady-state Navier-Stokes equations for hypersonic flow about blunt axisymmetric bodies

The steady-state Navier-Stokes equations are solved for hypersonic flow about blunt axisymmetric bodies. The equations of motion are solved by successive approximations using an implicit finite-difference scheme. The results are compared with viscous shock-layer theory, experimental data, and time-dependent solutions of the Navier-Stokes equations. It is demonstrated that viscous shock-layer theory is sufficiently accurate for the range of flight conditions normally encountered by entry vehicles.

Anderson, E. C.↗
NASA NTRS1976

Numerical solution of the steady-state Navier-Stokes equations for hypersonic flow about blunt axisymmetric bodies

The steady-state Navier-Stokes equations are solved for hypersonic flow about blunt axisymmetric bodies. The equations of motion are solved by successive approximations using an implicit finite-difference scheme. The results are compared with viscous shock-layer theory, experimental data, and time-dependent solutions of the Navier-Stokes equations. It is demonstrated that viscous shock-layer theory is sufficiently accurate for the range of flight conditions normally encountered by entry vehicles.

Anderson, E. C.↗
NASA NTRS1977

Low Reynolds number flow past a blunt axisymmetric body at angle of attack

Low Reynolds number flow of an ideal gas over a blunt axisymmetric body of large half-angle at small angles of attack is investigated, for the case of laminar hypersonic flow. Time-varying viscous shock layer equations describing the flowfield are obtained from the full Navier-Stokes system by keeping terms to second order in the inverse square root of Re in both viscous and inviscid regions; the equations are valid for moderate to high Re. Drag, skin friction, and heating rates were obtained at small (or zero) angles of attack. Conditions experienced by planetary entry probes during the high-altitude (early) legs of an atmospheric entry trajectory are pertinent to the problem.

Kumar, A.↗
NASA NTRS1991

An approximate viscous shock layer technique for calculating chemically reacting hypersonic flows about blunt-nosed bodies

An approximate axisymmetric method was developed which can reliably calculate fully viscous hypersonic flows over blunt nosed bodies. By substituting Maslen's second order pressure expression for the normal momentum equation, a simplified form of the viscous shock layer (VSL) equations is obtained. This approach can solve both the subsonic and supersonic regions of the shock layer without a starting solution for the shock shape. The approach is applicable to perfect gas, equilibrium, and nonequilibrium flowfields. Since the method is fully viscous, the problems associated with a boundary layer solution with an inviscid layer solution are avoided. This procedure is significantly faster than the parabolized Navier-Stokes (PNS) or VSL solvers and would be useful in a preliminary design environment. Problems associated with a previously developed approximate VSL technique are addressed before extending the method to nonequilibrium calculations. Perfect gas (laminar and turbulent), equilibrium, and nonequilibrium solutions were generated for airflows over several analytic body shapes. Surface heat transfer, skin friction, and pressure predictions are comparable to VSL results. In addition, computed heating rates are in good agreement with experimental data. The present technique generates its own shock shape as part of its solution, and therefore could be used to provide more accurate initial shock shapes for higher order procedures which require starting solutions.

Cheatwood, F. Mcneil↗