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Dejarnette, Fred R.

Publications and source records attributed to Dejarnette, Fred R..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

An approximate method for calculating heating rates on three-dimensional vehicles

An approximate method for calculating heating rates on three-dimensional vehicles at angle of attack is presented. The method is based on the axisymmetric analog for three-dimensional boundary layers and uses a generalized body fitted coordinate system. Edge conditions for the boundary layer solution are obtained from an inviscid flowfield solution, and because of the coordinate system used the method is applicable to any blunt body geometry for which a inviscid flowfield solution can be obtained. It is validated by comparing with experimental heating data and with Navier-Stokes calculations on the Shuttle orbiter at both wind tunnel and flight conditions and with Navier-Stokes calculations on the HL-20 at wind tunnel conditions.

Hamilton, H. H., II↗
NASA NTRS1993

Design issues for Mars planetary rovers

The paper presents some of the design issues and vehicle requirements that need to be addressed for the Mars planetary rovers. Some of the designs currently being investigated, including the JPL rover, the Martin Marietta vehicle, and the French Space Agency's VAP project, are examined. The rover must satisfy such mission requirements as surveying the terrain, preparing the landing sites, loading and unloading components for base operations, and aiding in the recovery of in situ materials.

Lee, Gordon K. F.↗
NASA NTRS1992

Boundary layer study on nozzle wall at hypersonic velocities

The boundary layer on the wall of the Hypervelocity Tunnel 9 was investigated with pitot pressure and total temperature measurements. Experimental results are presented for standard and supercooled Mach 14 runs. The boundary layer data at supercooled conditions are compared to numerical predictions made with a Navier-Stokes algorithm including vibrational nonequilibrium and intermolecular force effects. For standard tunnel conditions, the numerical solutions agree well with experimental data. For the supercooled cases, the numerical code predicts the total temperature but overpredicts the pitot pressure.

Jones, Kenneth M.↗
NASA NTRS1992

Wall interference calculation in a transonic test section including simulation of discrete slots

A computational simulation of a transonic wind tunnel test section with longitudinally slotted walls is described. This nonlinear slot model includes dynamic-pressure effects and a plenum pressure constraint. The simulation method developed is found to be a useful tool for analyzing the nature of the flowfield that exists in a longitudinally slotted transonic test section. Results obtained from the discrete-slot model are similar to those of a homogeneous model when a plenum pressure coefficient of zero is utilized, indicating little effect of slot discreteness with the transport aircraft models examined.

Al-Saadi, Jassim A.↗
NASA NTRS1992

Mars exploration advances: Missions to Mars - Mars base

An overview is presented of Mars missions and related planning with attention given to four mission architectures in the light of significant limitations. Planned unpiloted missions are discussed including the Mars Orbital Mapping Mission, the Mars Rover Sample Return, the Mars Aeronomy Orbiter, and the Mars Environmental Survey. General features relevant to the missions are mentioned including launch opportunities, manned-mission phases, and propulsion options. The four mission architectures are set forth and are made up of: (1) the Mars-exploration infrastructures; (2) science emphasis for the moon and Mars; (3) the moon to stay and Mars exploration; and (4) space resource utilization. The possibility of robotic missions to the moon and Mars is touched upon and are concluded to be possible by the end of the century. The ramifications of a Mars base are discussed with specific reference to habitability and base activities, and the human missions are shown to require a heavy-lift launcher and either chemical/aerobrake or nuclear-thermal propulsion system.

Dejarnette, Fred R.↗
NASA NTRS1992

An engineering aerodynamic heating method for hypersonic flow

A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.

Riley, Christopher J.↗
NASA NTRS1992

Approximate two layer (inviscid/viscous) methods to model aerothermodynamic environments

Approximate inviscid and boundary layer techniques for aerodynamic heating calculations are discussed. An inviscid flowfield solution is needed to provide surface pressures and boundary-layer edge properties. Modified Newtonian pressures coupled with an approximate shock shape will suffice for relatively simple shapes like sphere-cones with cone half-angles between 15 and 45 deg. More accurate approximate methods have been developed which make use of modified Maslen techniques. Slender and large angle sphere-cones and more complex shapes generally require an Euler code, like HALIS, to provide that information. The boundary-layer solution is reduced significantly by using the axisymmetric analog and approximate heating relations developed by Zoby, et al. (1981). Analysis is presented for the calculation of inviscid surface streamlines and metrics. Entropy-layer swallowing effects require coupling the inviscid and boundary-layer solutions.

Dejarnette, Fred R.↗
NASA NTRS1991

Effect of nose shape on three-dimensional stagnation region streamlines and heating rates

A new method for calculating the three-dimensional inviscid surface streamlines and streamline metrics using Cartesian coordinates and time as the independent variable of integration has been developed. The technique calculates the streamline from a specified point on the body to a point near the stagnation point by using a prescribed pressure distribution in the Euler equations. The differential equations, which are singular at the stagnation point, are of the two point boundary value problem type. Laminar heating rates are calculated using the axisymmetric analog concept for three-dimensional boundary layers and approximate solutions to the axisymmetric boundary layer equations. Results for elliptic conic forebody geometries show that location of the point of maximum heating depends on the type of conic in the plane of symmetry and the angle of attack, and that this location is in general different from the stagnation point. The new method was found to give smooth predictions of heat transfer in the nose region where previous methods gave oscillatory results.

Hassan, Basil↗
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↗
NASA NTRS1991

Engineering calculations of three-dimensional inviscid hypersonic flow fields

An approximate solution technique has been developed for three-dimensional, inviscid, hypersonic flows. The method uses Maslen's explicit pressure equation and the assumption of approximate stream surfaces in the shock layer. This approximation represents a simplification of Maslen's asymmetric method. The solution procedure involves iteratively changing the shock shape in the subsonic-transonic region until the correct body shape is obtained. Beyond this region, the shock surface is determined by using a marching procedure. Results are presented herein for a paraboloid and elliptic cone at angle of attack. Calculated surface pressure distributions, shock shapes, and property profiles are compared with experimental data and finite-difference solutions of the Euler equations. Comparisons of the results of the present method with experimental data and detailed predictions are very good. Since the present method provides a very rapid computational procedure, it can be used for parametric or preliminary design applications. One useful application would be to incorporate a heating procedure for aerothermal studies.

Riley, Christopher J.↗
NASA NTRS1990

An approximate method for calculating three-dimensional inviscid hypersonic flow fields

An approximate solution technique was developed for 3-D inviscid, hypersonic flows. The method employs Maslen's explicit pressure equation in addition to the assumption of approximate stream surfaces in the shock layer. This approximation represents a simplification to Maslen's asymmetric method. The present method presents a tractable procedure for computing the inviscid flow over 3-D surfaces at angle of attack. The solution procedure involves iteratively changing the shock shape in the subsonic-transonic region until the correct body shape is obtained. Beyond this region, the shock surface is determined using a marching procedure. Results are presented for a spherically blunted cone, paraboloid, and elliptic cone at angle of attack. The calculated surface pressures are compared with experimental data and finite difference solutions of the Euler equations. Shock shapes and profiles of pressure are also examined. Comparisons indicate the method adequately predicts shock layer properties on blunt bodies in hypersonic flow. The speed of the calculations makes the procedure attractive for engineering design applications.

Riley, Christopher J.↗