Search NASA⌕ Search

Engineering topics

Singh, D. J.

Publications and source records attributed to Singh, D. J..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

Thrust Enhancement in Hypervelocity Nozzles by Chemical Catalysis

In the hypersonic flight regime, the air-breathing supersonic combustion ramjet (scramjet) has been shown to be a viable propulsion system. The current designs of scramjet engines provide performance benefits only up to a Mach number of 14. Performance losses increase rapidly as the Mach number increases. To extend the applicability of scram'jets beyond Mach 14, research is being conducted in the area of inlet and wave drag reduction, skin-friction and heat-transfer reduction, nozzle loss minimization, low-loss mixing, and combustion enhancement. For high Mach number applications, hydrogen is the obvious fuel choice because of its high energy content per unit mass in comparison with conventional fuels. These flight conditions require engines to operate at supersonic internal velocities, high combustor temperatures, and low static pressures. The high static temperature condition enhances the production of radicals such as H and OH, and the low-pressure condition slows the reaction rates, particularly the recombination reactions. High-temperature and low-pressure constraints, in combination with a small residence time, result in a radical-rich exhaust gas mixture exiting the combustor. At high Mach number conditions (due to low residence time), H and OH do not have enough time to recombine ; thus, a significant amount of energy is lost as these high-energy free radical are exhausted. The objective of the present study is to conduct a flowfield analysis for a typical nozzle geometry for NASP-type vehicle to assess for thrust enhancement in hypervelocity nozzles by substituting small amount of phosphine for hydrogen.

Singh, D. J.↗
NASA NTRS1994

Numerical simulation of shock-induced combustion past blunt bodies using shock-fitting technique

Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock-induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A finite-difference, shock-fitting method is used to solve the complete set of Navier-Stokes and species conservation equations. In this approach, the bow shock represents a boundary of the computational domain and is treated as a discontinuity across which Rankine-Hugoniot conditions are applied. All interior details of the flow such as compression waves, reaction front, and the wall boundary layer are captured automatically in the solution. Since shock-fitting approach reduces the amount of artificial dissipation, all the intricate details of the flow are captured much more clearly than has been possible with the shock-capturing approach. This has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one-dimensional wave-interaction model than before.

Ahuja, J. K.↗
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 NTRS1994

Quasiglobal reaction model for ethylene combustion

The objective of this study is to develop a reduced mechanism for ethylene oxidation. The authors are interested in a model with a minimum number of species and reactions that still models the chemistry with reasonable accuracy for the expected combustor conditions. The model will be validated by comparing the results to those calculated with a detailed kinetic model that has been validated against the experimental data.

Singh, D. J.↗
NASA NTRS1994

Numerical Simulation of Shock-Induced Combustion Past Blunt Bodies Using Shock-Fitting Technique

Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A finite-difference, shock-fitting method is used to solve the complete set of Navier Stokes and species conservation equations. In this approach, the bow shock represents a boundary of the computational domain and is treated as a discontinuity across which Rankine-Hugoniot conditions are applied. All interior details of the flow such as compression waves, reaction front, and the wall boundary layer are captured automatically in the solution. Since shock-fitting approach reduces the amount of artificial dissipation, all the intricate details of the flow are captured much more clearly than has been possible with the shock-capturing approach. This has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one dimensional wave-interaction model than before.

Ahuja, J. K.↗
NASA NTRS1993

Numerical study of the performance of swept, curved compression surface scramjet inlets

The paper describes a computational performance-enhancement study in which systematic modifications were applied to a planar-sidewall compression scramjet inlet operating at an entrance Mach number of 4 and at a dynamic pressure of 2040 lbs/sq ft, and the performance of the modified inlet was compared to that of the original geometry by evaluating numerical flowfield solutions. The planar-sidewall compression angle was modified as a function of height using sidewall curvature and simultaneously employing both forward-swept and reverse-swept compression surfaces. The results of calculations illustrate the feasibility of obtaining enhanced performance via contour modifications of a baseline swept-sidewall compression scramjet inlet operating at an entrance Mach number of 4.

Korte, John J.↗
NASA NTRS1993

A dynamic grid adaptation procedure for mixing and reacting processes in scramjet combustors with injection

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 and boundary layers. Numerical studies of perpendicular sonic N2 injection and mixing in a Mach-3.8 stream and of the mixing and reaction effects from normal injection of H2 in a Mach-2.5 scramjet combustor environment are discussed. A dynamic grid adaptation procedure based on equilibration of spring-mass systems is employed to enhance the description of the complicated flow features. Numerical results are compared with experimental measurements, and conclusions are drawn in regard to the capability of the modeling procedure to describe the flow features associated with scramjet combustor components.

Ramakrishnan, R.↗
NASA NTRS1992

Investigation of hypersonic shock-induced combustion in a hydrogen-air system

A numerical study is conducted to simulate the ballistic range experiments at Mach 5.11 and 6.46. The flow field is found to be unsteady with periodic instabilities originating in the stagnation zone. The unsteadiness of the flow field decreased with increase in the Mach number, thus indicating that it is possible to stabilize such flow fields with a high degree of overdrive. The frequency of periodic instability is determined using Fourier power spectrum and is found to be in good agreement with the experimental data. The physics of the instability is explained by the wave interaction models available in the literature.

Ahuja, J. K.↗
NASA NTRS1992

Mixing and combustion effects in a sliding-wedge ram accelerator with hydrogen injection

The ram accelerator concept has been envisioned as the basis of a ground-based flight-test facility concept in which the test section follows the ram-accelerated projectile, and allows hydrogen mixing/combustion experiments to be conducted when hydrogen carried by the projectile is injected into the freestream. A numerical simulation is presently conducted for such mixing and combustion; the chemical reactions in question are modeled using a seven-step, seven-species model. A grid-adaptation procedure is used to resolve flow features in areas of high fluid and species gradients.

Ramakrishnan, R.↗
NASA NTRS1992

Premixed shock-induced combustion studies in the hypulse facility

Engine performance calculations strongly suggest that the oblique detonation-wave engine can outperform the diffusive-burning scramjet engine in the high hypervelocity flight regime; i.e., at flight Mach numbers exceeding about 14. Research is described that examines and characterizes the behavior of oblique detonation waves (ODW). The effort is both computational and experimental, the latter utilizing the NASA Hypulse expansion tube. The experimental work is hampered by preignition of the hydrogen-oxygen-helium mixtures in either the intermediate or acceleration sections of the facility. The experiments are described and possible causes of this preignition are discussed. Initial results obtained from a parallel computation effort indicate that temporally steady ODW are achievable in the Hypulse facility.

Chinitz, W.↗
NASA NTRS1992

CAN-DO, CFD-based Aerodynamic Nozzle Design and Optimization program for supersonic/hypersonic wind tunnels

A design program is developed which incorporates a modern approach to the design of supersonic/hypersonic wind-tunnel nozzles. The approach is obtained by the coupling of computational fluid dynamics (CFD) with design optimization. The program can be used to design a 2D or axisymmetric, supersonic or hypersonic, wind-tunnel nozzles that can be modeled with a calorically perfect gas. The nozzle design is obtained by solving a nonlinear least-squares optimization problem (LSOP). The LSOP is solved using an iterative procedure which requires intermediate flowfield solutions. The nozzle flowfield is simulated by solving the Navier-Stokes equations for the subsonic and transonic flow regions and the parabolized Navier-Stokes equations for the supersonic flow regions. The advantages of this method are that the design is based on the solution of the viscous equations eliminating the need to make separate corrections to a design contour, and the flexibility of applying the procedure to different types of nozzle design problems.

Korte, John J.↗
NASA NTRS1992

Three-dimensional simulation of a translating strut inlet

A three-dimensional Navier-Stokes code is used to numerically simulate the flow through a translating strut scramjet inlet. The inlet has variable geometry for efficient operation over a wide speed range. Overall flow-field features such as the corner flow, topwall separation, shockwave coalescence, cowl pressure increase, and flow distortion at the throat are investigated. Comparisons are made with experimental results to provide for the assessment of the present analysis. Effects of boundary-layer ingestion on the overall flow features are also investigated.

Singh, D. J.↗
NASA NTRS1992

Effects of nose bluntness and angle of attack on slender bodies in hypersonic flows

The effects of angle of attack and nose bluntness on the flow field and wall quantities are investigated for hypersonic flows of air over slender bodies. The bodies considered are slender cones and straight biconic configurations. The numerical procedures used are based on the solution of complete Navier-Stokes equations in the nose region and parabolized Navier-Stokes equations in the downstream region. Results are obtained for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The postshock flow field is studied in detail from the contour plots of Mach number, density, and temperature. Flow separation is observed on the leeward plane for an on-axis, 12.84 deg/7 deg (fore-cone and aft-cone angles) biconic geometry at 12 deg angle of attack. Also, the windward and leeward heating rates for the fore-cone section decrease by a factor of four and five, respectively, when the nose bluntness is increased by an order of magnitude. The effect of nose bluntness for slender cone persists as far as 200 nose radii downstream.

Tiwari, S. N.↗
NASA NTRS1992

Combined effect of nose bluntness and angle of attack on slender bodies in viscous hypersonic flows

Hypersonic flows over cones and straight biconic configurations are calculated for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. Effect of angle of attack and nose bluntness on these slender cones in air is studied extensively. The numerical procedures are based on the solution of complete Navier-Stokes equations at the nose section and parabolized Navier-Stokes equations further downstream. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The complete flow field is thoroughly analyzed with respect to velocity, temperature, pressure, and entropy profiles. The post shock flow field is studied in detail from the contour plots of Mach number, density, pressure, and temperature. The effect of nose bluntness for slender cones persists as far as 200 nose radii downstream.

Tiwari, S. N.↗
NASA NTRS1992

Numerical simulation of shock-induced combustion/detonation

A numerical study is conducted to simulate the shock-induced combustion in premixed H2-air mixtures. Two types of bodies, blunt (spherical projectile) and sharp (wedge), are considered in the study. A nine-species, 18-step finite-rate H2-air chemical reaction mechanism coupled with the Navier-Stokes equations is solved. The flow field over the blunt body is found to be unsteady, when the projectile velocity is same as the Chapman-Jouget velocity of the mixture. The unsteadiness is caused by the periodic instabilities originating in the stagnation zone. Numerical results show good qualitative agreement with the ballistic range shadowgraph. In addition, the frequency of oscillations, determined by using the Fourier power spectrum, is found to be in good agreement with the experiment. The flow field over the wedge is found to be stable for the conditions considered in this study. The oblique detonation wave structure is investigated and the important flow features are discussed.

Singh, D. J.↗
NASA NTRS1991

Least-squares/parabolized Navier-Stokes procedure for optimizing hypersonic wind tunnel nozzles

A new procedure is demonstrated for optimizing hypersonic wind-tunnel-nozzle contours. The procedure couples a CFD computer code to an optimization algorithm, and is applied to both conical and contoured hypersonic nozzles for the purpose of determining an optimal set of parameters to describe the surface geometry. A design-objective function is specified based on the deviation from the desired test-section flow-field conditions. The objective function is minimized by optimizing the parameters used to describe the nozzle contour based on the solution to a nonlinear least-squares problem. The effect of the changes in the nozzle wall parameters are evaluated by computing the nozzle flow using the parabolized Navier-Stokes equations. The advantage of the new procedure is that it directly takes into account the displacement effect of the boundary layer on the wall contour. The new procedure provides a method for optimizing hypersonic nozzles of high Mach numbers which have been designed by classical procedures, but are shown to produce poor flow quality due to the large boundary layers present in the test section. The procedure is demonstrated by finding the optimum design parameters for a Mach 10 conical nozzle and a Mach 6 and a Mach 15 contoured nozzle.

Korte, John J.↗