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NASA NTRS1993

Summary of the GASP code application and evaluation effort for scramjet combustor flowfields

Numerical simulations of 3D turbulent mixing and reacting flows have been systematically evaluated by comparison with experimental data and other numerical solutions to provide confidence in the General Aerodynamic Simulation Program (GASP) code version 1.3. Specifically, the GASP flow solver has been used to model high speed flow through scram jet combustors. Unit injector problems evaluated include (1) UVA staged sonic normal injection of air into Mach 2 air stream, (2) the VPI Mach 1.7 low angled (15 deg) flush wall helium injection into an unconfined Mach 6 'cold air' stream, and (3) the HYPULSE angled (30 deg) flush wall hydrogen injection into a Mach 6, high enthalpy (flight Mach 17 simulation) nitrogen and air test gases. Results of the numerical simulation for the first two test cases have been compared with detailed in-stream measurements and with SPARK CFD solutions. The results for the last two cases have been compared with measured wall pressure and heat flux data and with SPARK CFD solutions. The present GASP solutions compare favorably with both the experimental data and the SPARK solutions.

Srinivasan, Shivakumar↗

NASA NTRSDate not supplied

Sonic Boom Ground Noise Minimization Via the Adjoint Method

A multidisciplinary design optimization methodology to directly minimize the ground-level noise generated by the sonic boom of a high altitude supersonic body is presented. A Cartesian Euler flow solver is coupled with an atmospheric propagation tool to create a ground-noise analysis capability for supersonic bodies. Adjoint formulations for both the flow solver and propagation tool are also coupled to compute sensitivities of shape variations in the body in a highly efficient manner. A gradient-based optimizer is then introduced to forge a valuable design capability. Output-based mesh adaptation that is driven directly by ground-level noise is employed to increase accuracy and provide error estimation. The design method is demonstrated first on a simple axisymmetric body with few design variables to evaluate the efficacy of the optimization scheme. Guided by the results of this initial case, the problem is then repeated with somewhat different design variables to further demonstrate the capabilities of the design method. Finally, the method is applied to a real-world problem by optimizing control surface deflection settings of a low-boom aircraft to minimize ground noise while maintaining trimmed, level flight.

ARMD↗

NASA NTRS1977

Sonic boom analysis for high-altitude flight at high Mach number

Numerical programs for the computation of the flow field from the airplane at the flight altitude to the ground are presented. They take into account the nonlinear effects of high Mach number, the entropy change across the shock, the entropy and enthalpy variations in the atmospheric layer, and the gravitational effect. Extension of the programs for the axisymmetric problems to handle nonaxisymmetric terms is described. The asymmetry can be caused by the geometry of the body and the lift, and also by the fact that the variations in the atmospheric layer are two-dimensional. Numerical results for ground level signatures of several configurations at various flight conditions are presented and compared with existing approximate theories to demonstrate the influences of these nonlinear effects.

Ferri, A.↗

NASA NTRS1979

Numerical solution for perpendicular sonic hydrogen injection into a ducted supersonic airstream

This note discusses a computer program being developed to study the flow field near opposing perpendicular fuel injectors in scramjets. The MacCormack time-split, finite difference relaxation technique was used to solve the full two-dimensional compressible Navier-Stokes equations along with energy and species equations. By using this technique, a program was developed to consider the turbulent nonreacting flow of hydrogen and air in a rectangular duct. A damping term, proportional to the second derivative of pressure and temperature, was used to produce a stable solution behind the hydrogen jet in the neighborhood of the recompression shock. A case using actual conditions encountered in current scramjet design was analyzed, with results agreeing qualitatively with experimental observations.

Drummond, J. P.↗

NASA NTRS1954

The calculation of pressure on slender airplanes in subsonic and supersonic flow

Under the assumption that a wing, body, or wing-body combination is slender or flying at near sonic velocity, expressions are given which permit the calculation of pressure in the immediate vicinity of the configuration. The disturbance field, in both subsonic and supersonic flight, is shown to consist of two-dimensional disturbance fields extending laterally and a longitudinal field that depends on the streamwise growth of cross-sectional area. A discussion is also given of couplings, between lifting and thickness effects, that necessarily arise as a result of the quadratic dependence of pressure on the induced velocity components. (author)

Heaslet, Max A↗

NASA NTRS1994

Modeling scramjet combustor flowfields with a grid adaptation scheme

The accurate description of flow features associated with the normal injection of fuel into supersonic primary flows is essential in the design of efficient engines for hypervelocity aerospace vehicles. The flow features in such injections are complex with multiple interactions between shocks and between shocks boundary layers. Numerical studies of perpendicular sonic N2 injection and mixing in a Mach 3.8 scramjet combustor environment are discussed. A dynamic grid adaptation procedure based on the equilibration of spring-mass system is employed to enhanced the description of the complicated flow features. Numerical results are compared with experimental measurements and indicate that the adaptation procedure enhances the capability of the modeling procedure to describe the flow features associated with scramjet combustor components.

Ramakrishnan, R.↗

NASA NTRS1973

A method for predicting shock shapes and pressure distributions on two dimensional airfoils at large angles of attack

A method is presented for determining shock envelopes and pressure distributions for two-dimensional airfoils at angles of attack sufficiently large to cause shock detachment and subsonic flow over the windward surface of the airfoil. Correlation functions obtained from exact solutions are used to relate the shock standoff distance at the stagnation and sonic points of the body through a suitable choice for the shock shape. The necessary correlation functions were obtained from perfect gas solutions but may be extended to any gas flow for which the normal shock-density ratio can be specified.

Kaattari, G. E.↗

NASA NTRS1992

Sonic boom prediction for the Langley Mach 2 low-boom configuration

Sonic boom pressure signatures and aerodynamic force data for the Langley Mach 2 low sonic boom configuration were computed using the TranAir full-potential code. A solution-adaptive Cartesian grid scheme is utilized to compute off-body flow field data. Computations were performed with and without nacelles at several angles of attack. Force and moment data were computed to measure nacelle effects on the aerodynamic characteristics and sonic boom footprints of the model. Pressure signatures were computed both on and off ground-track. Near-field pressure signature computations on ground-track were in good agreement with experimental data. Computed off ground-track signatures showed that maximum pressure peaks were located off ground-track and were significantly higher than the signatures on ground-track. Bow shocks from the nacelle inlets increased lift and drag, and also increased the magnitude of the maximum pressure both on and off ground-track.

Madson, Michael D.↗

NASA NTRS1992

A three-dimensional axisymmetric photochemical flow model of the cometary 'inner' shock layer

Assuming the Newtonian thin layer approximation to describe the structure of the shock layer between the cometary 'ionopause' and the inner shock, a 3D axisymmetric photochemical flow model of this layer is constructed. While sources of ions in this layer are the flow across the inner shock and photoionization of neutrals within it, the sinks are the flow into the flanks and dissociative recombination, the latter being the dominant one. For Halley's comet at the time of the Giotto encounter, the calculated thickness of the layer is very small, typically about 100 km for expected values of the dissociative-recombination coefficient. This is not inconsistent with the observations. The lateral flow speed near the point of encounter (inbound) is about 0.9 km/s, while the sonic line is at an angle of about 50 deg to the sun-comet line. Testing the validity of this model will have to await a cometary rendezvous mission such as the proposed CRAF/Cassini mission.

Damas, M. C.↗

NASA NTRS1998

Mixing of Supersonic Streams

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane. Initial cold flow testing of the model is underway to determine both, the behavior of the ingested air in the duct and to validate the mixing diagnostics. During the tests, each of the two rocket nozzles ejected up to two pounds mass per second into the 13.6 square inch duct. The tests showed that the mass flow of the rockets was great enough to cause the entrained air to go sonic at the strut, which is the location of the rocket nozzles. More tests are necessary to determine whether the entrained air chokes due to the reduction in the area of the duct at the strut (a physical choke), or because of the addition of mass inside the duct at the nozzle exit (a Fabri choke). The initial tests of the mixing diagnostics are showing promise.

Hawk, Clark W.↗

NASA NTRS1992

Ground signature extrapolation of three-dimensional near-field CFD predictions for several HSCT configurations

A CFD analysis of the near-field sonic boom environment of several low boom High Speed Civilian Transport (HSCT) concepts is presented. The CFD method utilizes a multi-block Euler marching code within the context of an innovative mesh topology that allows for the resolution of shock waves several body lengths from the aircraft. Three-dimensional pressure footprints at one body length below three-different low boom aircraft concepts are presented. Models of two concepts designed by NASA to cruise at Mach 2 and Mach 3 were built and tested in the wind tunnel. The third concept was designed by Boeing to cruise at Mach 1.7. Centerline and sideline samples of these footprints are then extrapolated to the ground using a linear waveform parameter method to estimate the ground signatures or sonic boom ground overpressure levels. The Mach 2 concept achieved its centerline design signature but indicated higher sideline booms due to the outboard wing crank of the configuration. Nacelles are also included on two of NASA's low boom concepts. Computations are carried out for both flow-through nacelles and nacelles with engine exhaust simulation. The flow-through nacelles with the assumption of zero spillage and zero inlet lip radius showed very little effect on the sonic boom signatures. On the other hand, it was shown that the engine exhaust plumes can have an effect on the levels of overpressure reaching the ground depending on the engine operating conditions. The results of this study indicate that engine integration into a low boom design should be given some attention.

Siclari, M. J.↗

NASA NTRS2008

Output-Adaptive Tetrahedral Cut-Cell Validation for Sonic Boom Prediction

A cut-cell approach to Computational Fluid Dynamics (CFD) that utilizes the median dual of a tetrahedral background grid is described. The discrete adjoint is also calculated, which permits adaptation based on improving the calculation of a specified output (off-body pressure signature) in supersonic inviscid flow. These predicted signatures are compared to wind tunnel measurements on and off the configuration centerline 10 body lengths below the model to validate the method for sonic boom prediction. Accurate mid-field sonic boom pressure signatures are calculated with the Euler equations without the use of hybrid grid or signature propagation methods. Highly-refined, shock-aligned anisotropic grids were produced by this method from coarse isotropic grids created without prior knowledge of shock locations. A heuristic reconstruction limiter provided stable flow and adjoint solution schemes while producing similar signatures to Barth-Jespersen and Venkatakrishnan limiters. The use of cut-cells with an output-based adaptive scheme completely automated this accurate prediction capability after a triangular mesh is generated for the cut surface. This automation drastically reduces the manual intervention required by existing methods.

Park, Michael A.↗

NASA NTRS1990

Planar measurement of flow field parameters in a nonreacting supersonic combustor using laser-induced iodine fluorescence

A nonintrusive optical technique, laser-induced iodine fluorescence, has been used to obtain planar measurements of flow field parameters in the supersonic mixing flow field of a nonreacting supersonic combustor. The combustor design used in this work was configured with staged transverse sonic injection behind a rearward-facing step into a Mach 2.07 free stream. A set of spatially resolved measurements of temperature and injectant mole fraction has been generated. These measurements provide an extensive and accurate experimental data set required for the validation of computational fluid dynamic codes developed for the calculation of highly three-dimensional combustor flow fields.

Hartfield, Roy J., Jr.↗

NASA NTRS1976

A multi-modal solution for the transmission of sound in nonuniform ducts

The method of weighted residuals in the form of a modified Galerkin method with boundary residuals is developed for the study of the transmission of sound in nonuniform ducts carrying a steady compressible flow. In this formulation the steady flow is modeled as essentially one-dimensional but with a kinematic modification to force tangency of the flow at the duct walls. Good agreement has been obtained with known results for transmission and reflection coefficients in hard walled ducts up to near sonic velocities. In ducts with acoustically compliant boundaries good comparisons have been more difficult to achieve except at low Mach numbers. The problem of transmission in a straight, acoustically treated duct with a uniform flow has been formulated and the Galerkin method used with basis functions derived from the case when flow is absent. Results indicate that favorable comparisons with exact computations can be obtained if care is taken in choosing the basis functions.

Eversman, W.↗

NASA NTRS2015

Under-Track CFD-Based Shape Optimization for a Low-Boom Demonstrator Concept

The detailed outer mold line shaping of a Mach 1.6, demonstrator-sized low-boom concept is presented. Cruise trim is incorporated a priori as part of the shaping objective, using an equivalent-area-based approach. Design work is performed using a gradient-driven optimization framework that incorporates a three-dimensional, nonlinear flow solver, a parametric geometry modeler, and sensitivities derived using the adjoint method. The shaping effort is focused on reducing the under-track sonic boom level using an inverse design approach, while simultaneously satisfying the trim requirement. Conceptual-level geometric constraints are incorporated in the optimization process, including the internal layout of fuel tanks, landing gear, engine, and crew station. Details of the model parameterization and design process are documented for both flow-through and powered states, and the performance of these optimized vehicles presented in terms of inviscid L/D, trim state, pressures in the near-field and at the ground, and predicted sonic boom loudness.

Wintzer, Mathias↗

NASA NTRS1961

Pressure Loads Produced on a Flat-Plate Wing By Rocket Jets Exhausting in a Spanwise Direction Below the Wing and Perpendicular to a Free-Stream Flow of Mach Number 2.0

An investigation at a Reynolds number per foot of 14.4 x 10(exp 6) was made to determine the pressure loads produced on a flat-plate wing by rocket jets exhausting in a spanwise direction beneath the wing and perpendicular to a free-stream flow of Mach number 2.0. The ranges of the variables involved were (1) nozzle types - one sonic (jet Mach number of 1.00), two supersonic (jet Mach numbers of 1.74 and 3.04),. and one two-dimensional supersonic (jet Mach number of 1.71); (2) vertical nozzle positions beneath the wing of 4, 8 and 12 nozzle-throat diameters; and (3) ratios of rocket-chamber total pressure to free- stream static pressure from 0 to 130. The incremental normal force due to jet interference on the wing varied from one to two times the rocket thrust and generally decreased as the pressure ratio increased. The chordwise coordinate of the incremental-normal-force center of pressure remained upstream of the nozzle center line for the nozzle positions and pressure ratios of the investigation. The chordwise coordinate approached zero as the jet vertical distance beneath the wing increased. In the spanwise direction there was little change due to varying rocket-jet position and pressure ratio. Some boundary-layer flow separation on the wing was observed for the rocket jets close to the wing and at the higher pressure ratios. The magnitude of the chordwise and spanwise pressure distributions due to jet interference was greatest for rocket jets close to the wing and decreased as the jet was displaced farther from the wing. The design procedure for the rockets used is given in the appendix.

Falanga, Ralph A.↗

NASA NTRS1980

Comparison of experimental and computational shock structure in a transonic compressor rotor

Measurement of passage shock strength in a transonic compressor rotor using a gas fluorescent technique revealed an unexpected variation in shock strength in the radial direction. An axisymmetric idealization would normally predict that the passage shock strength would gradually weaken when moving radially inward until disappearing at the sonic radius. However, the measurements indicated a sharp peak in strength at the nominal sonic radius. Blade boundary layer separation originating at this point accounts for about one half of the total rotor losses. A numerical computation of the three-dimensional inviscid flow, using time-marching techniques, has accurately predicted in general the radial and tangential variations in passage shock strength and in particular the sharp pressure peak at the nominal sonic radius. The overall shock strength was somewhat overpredicted, but this overprediction may be the result of boundary layer separation in the experiment. This paper presents comparisons between the optical density measurements and computational results and in addition a short analytical discussion which demonstrates that the sharp shock strength rise may occur in many transonic compressor rotors.

Haymann-Haber, G.↗

NASA NTRS1976

Steady supersonic flowfields with embedded subsonic regions

Supersonic flow past a blunt body is considered, where the flow contains an embedded subsonic region which lies between the shock wave and the body surface and is bounded by sonic lines from the body to the shock. A numerical approach is taken, which uses a basic finite difference scheme that solves the unsteady fluid dynamic equations in integral form. The unsteady equations are everywhere hyperbolic in time so no distinction need be made between subsonic and supersonic regions. Solutions to the mixed elliptic and hyperbolic steady flow equations are approached asymptotically in time. The method is illustrated for two-dimensional flows.

Maccormack, R. W.↗