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Wieting, Allan R.

Publications and source records attributed to Wieting, Allan R..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Calibration of the Langley 8-Foot High Temperature Tunnel for Hypersonic Airbreathing Propulsion Testing

The NASA Langley 8-Foot High Temperature Tunnel has recently been modified to produce a unique testing capability for hypersonic airbreathing propulsion systems. Prior to these modifications, the facility was used primarily for aerothermal loads and structural verification testing at true flight total enthalpy conditions for Mach numbers between 6 and 7. One of the recent modifications was an oxygen replenishment system which allows operating airbreathing propulsion systems to be tested at true flight total enthalpies. Following the modifications to the facility, calibration runs were performed at total enthalpies corresponding to flight Mach numbers of 6.3 and 6.8 to establish the flow characteristics of the facility with its new capabilities. The results of this calibration, as well as modifications to tunnel combustor hardware prior to calibration to improve tunnel flow quality, are described in this paper.

Huebner, Lawrence D.↗
NASA NTRS1991

Multiple shock-shock interference on a cylindrical leading edge

The details of an experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet are presented. This Mach 8 study has provided the first detailed pressure and heat transfer rate distributions on a cylinder resulting from a two-dimensional shockwave interference pattern created by two incident oblique shock waves intersecting the cylinder bow shock wave. The peak heat transfer rate was 38 times the undisturbed flow stagnation point level and occurred when the two oblique shock waves coalesced prior to intersecting the cylinder bow shock wave. Development of pressure deflection diagrams identified a new interference pattern consisting of concomitant supersonic jets separated from each other by a shear layer and submerged in the subsonic region between the bow shock wave and body.

Wieting, Allan R.↗
NASA NTRS1991

Computational studies of a fluid spike as a leading edge protection device for shock-shock interference heating

The effectiveness of a fluid spike as a device to protect leading edges of hypersonic atmospheric flight vehicles from high aerothermal loads produced by complex shock-shock interference is studied. The two-dimensional Navier-Stokes equations are solved using an unstructured cell-centered, fully implicit, flux-difference split algorithm. Adaptively generated unstructured meshes are employed. A type IV shock-shock interference for Mach 8 flow on a cylindrical leading edge with and without a small contraflow supersonic jet (fluid spike) placed at two different locations on the body is solved. A typical flow past a blunt body with a type IV shock-shock interference produces very high pressures and heat fluxes on the leading edge. Present results indicate that a fluid spike displaces the bow shock further in front of the body and modifies the shock-shock interference pattern. This leads to reduced peak pressures and heat fluxes on the body.

Prabhu, Ramadas K.↗
NASA NTRS1990

Thermal-structural finite element analysis using linear flux formulation

A linear flux approach is developed for a finite element thermal-structural analysis of steady state thermal and structural problems. The element fluxes are assumed to vary linearly in the same form as the element unknown variables, and the finite element matrices are evaluated in closed form. Since numerical integration is avoided, significant computational time saving is achieved. Solution accuracy and computational speed improvements are demonstrated by solving several two and three dimensional thermal-structural examples.

Pandey, Ajay K.↗
NASA NTRS1990

Shock interference heating in scramjet engines

Experimental and analytical research sponsored by the NASA Langley Research center and the NASP Structures Technology Maturation Program to define critical aerothermal loads for the NASP engine is summarized. Presented is a review of (1) shock-shock interaction on the engine cowl leading edge that results in a supersonic jet impinging on the leading edge surface and causes the heat transfer rate to be amplified by a factor of 30 or more over the undisturbed (no shock interaction) flow stagnation point heat transfer rate, (2) the effectiveness of supersonic film cooling with and without the effects of an impinging oblique shock wave, and (3) oblique shock impingement in an axial compression corner.

Wieting, Allan R.↗
NASA NTRS1990

An adaptive refinement procedure for transient thermal analysis using nodeless variable finite elements

An adaptive mesh refinement procedure that uses nodeless variables and quadratic interpolation functions is presented for analyzing transient thermal problems. A temperature based finite element scheme with Crank-Nicolson time marching is used to obtain the thermal solution. The strategies used for mesh adaption, computing refinement indicators, and time marching are described. Examples in one and two dimensions are presented and comparisons are made with exact solutions. The effectiveness of this procedure for transient thermal analysis is reflected in good solution accuracy, reduction in number of elements used, and computational efficiency.

Ramakrishnan, R.↗
NASA NTRS1990

Shock/shock interference on a transpiration cooled hemispherical model

Experimental results are presented which show the effectiveness of transpiration cooling in reducing the peak heat flux caused by an impinging shock on a bow shock of a hemispherical model. The 12-inch diameter hemispherical transpiration model with helium coolant was tested in the Calspan 48-inch Hypersonic Shock Tunnel at nominal Mach 12.1 and freestream unit Reynolds number of 0.33 x 10 to the 6th/ft. An incident shock wave, generated by a blunt flat-plate shock generator inclined at 10 deg to the freestream, intersected the bow shock of the model to produce shock/shock interference. The stagnation heat flux without coolant or shock/shock interference was about 1.6 times a smooth surface laminar prediction due to effective roughness of the coolant ejection slots. A coolant mass flux 31 percent of the freestream mass flux reduced the stagnation heat flux to zero without shock/shock interference. However, for the same coolant mass flux and with shock/shock interference the peak heat flux was only reduced 8.3 percent, even though the total integrated heat load was reduced.

Nowak, Robert J.↗
NASA NTRS1990

Application of a finite element algorithm for high speed viscous flows using structured and unstructured meshes

A higher-order streamline upwinding Petrov-Galerkin finite element method is employed for high speed viscous flow analysis using structured and unstructured meshes. For a Mach 8.03 shock interference problem, successive mesh adaptation was performed using an adaptive remeshing method. Results from the finite element algorithm compare well with both experimental data and results from an upwind cell-centered method. Finite element results for a Mach 14.1 flow over a 24 degree compression corner compare well with experimental data and two other numerical algorithms for both structured and unstructured meshes.

Vemaganti, Gururaja R.↗
NASA NTRS1990

A structured and unstructured remeshing method for high speed flows

An adaptive remeshing method using both triangular and quadrilateral elements suitable for high speed flows is presented. For inviscid flows the method generates completely unstructured meshes. For viscous flows the boundary layer edge is identified adaptively and a structured mesh is generated in the boundary layer, and an unstructured mesh is generated in the inviscid region. Examples of inviscid and viscous mesh adaptations for high speed flows are presented. A comparison is made between first order and higher order finite element algorithms when used in association with the remeshing method.

Vemaganti, Gururaja R.↗
NASA NTRS1989

Finite element flow-thermal-structural analysis of aerodynamically heated leading edges

Hypersonic vehicles operate in a hostile aerothermal environment which has a significant impact on their aerothermostructural performance. Significant coupling occurs between the aerodynamic flow field, structural heat transfer, and structural response creating a multidisciplinary interaction. A long term goal of the Aerothermal Loads Branch at the NASA Langley Research Center is to develop a computational capability for integrated fluid, thermal and structural analysis of aerodynamically heated structures. The integrated analysis capability includes the coupling between the fluid and the structure which occurs primarily through the thermal response of the structure, because: (1) the surface temperature affects the external flow by changing the amount of energy absorbed by the structure, and (2) the temperature gradients in the structure result in structural deformations which alter the flow field and attendant surface pressures and heating rates. In the integrated analysis, a finite element method is used to solve: (1) the Navier-Stokes equations for the flow solution, (2) the energy equation of the structure for the temperature response, and (3) the equilibrium equations of the structure for the structural deformation and stresses.

Dechaumphai, Pramote↗
NASA NTRS1989

Effect of leading edge sweep on shock-shock interference at Mach 8

These Mach 8 experimental results are applicable to the details of a shock-shock interference that may occur on an engine inlet of a hypersonic vehicle from a swept forebody shock interacting with a swept cowl leading edge bow shock or from a swept splitter plate shock interacting with a swept fuel injection strut bow shock. Peak surface pressure and heat transfer rate on the cylinder were about 10 and 30 times the undisturbed flow stagnation point value, respectively, for the 0 deg sweep test. A comparison of the 15 deg and 30 deg swept results with the 0 deg swept results shows that peak pressure was reduced about 13 percent and 44 percent, respectively, and peak heat transfer rate was reduced about 7 percent and 27 percent, respectively.

Glass, Christopher E.↗
NASA NTRS1989

Thermal-structural finite element analysis using linear flux formulation

A linear flux approach is developed for a finite element thermal-structural analysis of steady-state thermal and structural problems. The element fluxes are assumed to vary linearly in the same form as the element unknown variables, and the finite element matrices are evaluated in closed form. Since numerical integration is avoided, significant computational time saving is achieved. Solution accuracy and computational speed improvements are demonstrated by solving several two-and three-dimensional thermal-structural examples.

Pandey, Ajay K.↗
NASA NTRS1989

Fluid-thermal-structural interaction of aerodynamically heated leading edges

A two-dimensional finite element approach is presented for the integrated fluid-thermal-structural analysis of aerodynamically heated leading edges. The approach is combined with an adaptive unstructured remeshing technique to solve the Navier-Stokes equations for high speed compressible flow, the energy equation for the structure thermal response, and the quasi-static equilibrium equations for the structural response. Coupling and interaction between the three disciplines are demonstrated using two applications for high speed flow over a cylinder and a simulated engine leading edge verification test.

Dechaumphai, Pramote↗