Engineering topics

Mcclinton, Charles R.

Publications and source records attributed to Mcclinton, Charles R..

15 records

NASA NTRS1992

Computation of losses in a scramjet combustor

The losses in a conceptual scramjet combustor at flight Mach numbers of 8, 10, 12, 16 and 20 are computed. These losses are extracted from three-dimensional parabolized Navier-Stokes solutions of the turbulent, reacting combustor flow field. A combustor performance index was defined based on the rationale that an efficient scramjet combustor should add heat to the fluid in such a manner as to maximize the stream thrust at the combustor exit while minimizing the losses. This index showed a decrease of more than 40 percent as the flight Mach number increased from 8 to 20, indicative of a drop in the thrust-producing potential of the scramjet at the upper end of the speed regime studied. A breakdown of the losses showed that dissipation, nonequilibrium chemistry and heat diffusion contributed roughly 15 percent, 35 percent, and 50 percent to the irreversible increase in entropy at Mach 8 and 22 percent, 13 and 65 percent at Mach 20.

Kamath, Pradeep S.↗

NASA NTRS1992

Scramjet combustor and nozzle computations

The SHIP3D PNS code is used to conduct a 3D parametric study of a scramjet combustor and nozzle configuration at flight Mach numbers of 8, 10, 12, 16, and 20. The study demonstrated the powerful computational strategy of using a highly efficient, modified PNS code and a nonkinetic chemistry model to analyze test data and conduct parametric studies of scramjet combustors and nozzles. The mixing efficiency was found to decrease by 0.2 when the flight Mach number increased from 8 to 12 and showed little change at higher Mach numbers. The protruding ramp injector generated an oblique shock that impinged on the cowl. Ramp-slot injection produced better mixing, both because of the interaction of this shock with the slot fuel, and because of better fuel distribution in the duct. The former effect was enhanced when the ramp angle was increased. Mixing was also enhanced when the ramp fuel injection angle and sweep angle were increased.

Kamath, Pradeep S.↗

NASA NTRS1991

Scramjet combustor analysis with the SHIP3D PNS code

The SHIP3D PNS code has the capability to perform 3D computations of scramjet combustors efficiently at flight Mach numbers of 10 and above, where the combustor flow is mostly supersonic. The fast solution times make skin friction and heat transfer calculations feasible and permit the utilization of multiequation turbulence models. Application of this relatively uncomplicated methodology to compute scramjet combustor flow fields and heat transfer is presented with particular emphasis on the treatment of ramp fuel injectors.

Kamath, Pradep S.↗

NASA NTRS1991

CFD evaluation of Mach 17 HYPULSE scramjet combustor data

Three-dimensional finite rate chemistry solutions are performed on a single fuel injector configuration. The results are compared with limited experimental data obtained from the HYPULSE expansion tube facility at simulated flight Mach 17 flow conditions. All comparisons, except for wall heat flux, were in excellent agreement. Key findings from this study are useful in interpretation of the experimental measurements.

Bobskill, Glenn J.↗

NASA NTRS1991

A parametric study of scramjet combustors for Mach 8 to Mach 20 flight

Performance trends for a combustor with ramp and slot fuel injectors are investigated through a three-dimensional parametric study at flight Mach numbers of 8, 10, 12, 16, and 20. Focus is placed on the mixing (combustion) efficiency, combustor loads, and thrust over the Mach number range. It is observed that the mixing efficiency decreases by 0.2 when a flight Mach number is increased from 8 to 12 but changes little for higher Mach numbers; the lower mixing is attributed more to a large decrease in the fuel-to-air velocity ratio and residence time over the Mach number range than to compressibility. It is pointed out that the ramp generates a shock impinging on the cowl, while a combined ramp and slot injection produces better mixing, both because of the action of this shock on the slot fuel and because of better fuel distribution in the duct.

Kamath, Pradeep S.↗

NASA NTRS1991

Numerical study of external burning flowfields

This paper demonstrates the successful application of CFD to modeling an external burning flowfield. The study used the 2D, 3D, and PNS versions of the SPARK code. Various grids, boundary conditions, and ignition methodologies have been employed. Flameholding was achieved through the use of a subsonic outflow condition and a hot block located behind the step to ignite the fuel. Since the resulting burning produces a large subsonic region downstream of the cowl, this entire surface can be pressurized to the level of the back pressure. An evaluation of interactions between the ramjet exhaust and the external burning products demonstrate the complexity of this design issue. Ths code is now capable of evaluating the external burning effectiveness for flight vehicles using simple injector schemes, and the methodology can be readily applied to other external burning designs.

Bittner, Robert D.↗

NASA NTRS1991

Numerical simulation of transverse fuel injection

A review of recent work at NASA Langley Research Center to compare the predictions of transverse fuel injector flow fields and mixing performance with experimental results is presented. Various cold (non-reactive) mixing studies were selected for code calibration which include the effects of boundary layer thickness and injection angle for sonic hydrogen injection into supersonic air. Angled injection of helium is also included. This study was performed using both the three-dimensional elliptic and the parabolized Navier-Stokes (PNS) versions of SPARK. Axial solution planes were passed from PNS to elliptic and elliptic to PNS in order to efficiently generate solutions. The PNS version is used both upstream and far downstream of the injector where the flow can be considered parabolic in nature. The comparisons are used to identify experimental deficiencies and computational procedures to improve agreement.

Mao, Marlon↗

NASA NTRS1991

A highly efficient engineering tool for three-dimensional scramjet flowfield and heat transfer computations

The SIMPLE-based parabolic flow code, SHIP3D, was under development for use as a parametric design and analysis tool for scramjets. Some capabilities and applications of the code are demonstrated, and a report on its current status is given. The focus is on the combustor for which the code was mostly used. Recently, it was also applied to nozzle flows. Code validation results are presented for combustor unit problems involving film cooling, transverse fuel injection, and nozzle test. A parametric study of a film cooled or transpiration cooled Mach 16 combustor is also conducted to illustrate the application of the code to a design problem.

Kamath, Pradeep S.↗

NASA NTRS1990

CFD support of NASP design

This paper presents a summary of design studies from the 'open' literature which illustrate the level of effort and the use of computational fluid dynamics (CFD) to support the National Aerospace Plane (NASP) X-30 design. CFD plays a major role in the NASP program, particularly for the very high speed regions (Mach greater than 10) where wind tunnels cannot fully simulate the flow, and flow field measurements are difficult to obtain. Full simulation (nose-to-tail analysis) of the NASP flow field, both internal and external, is discussed.

Mcclinton, Charles R.↗

NASA NTRS1990

CFD support of NASP design

This paper presents a summary of design studies from the 'open' literature which illustrate the level of effort and the use of computational fluid dynamics (CFD) to support the National Aerospace Plane (NASP) X-30 design. CFD plays a major role in the NASP program, particularly for the very high speed regions (Mach greater than 10) where wind tunnels cannot fully simulate the flow and/or flow-field measurements are difficult to obtain. Full simulation (nose-to-tail analysis) of the NASP flow field, both internal and external, is discussed.

Mcclinton, Charles R.↗

NASA NTRS1990

A computational investigation of flow losses in a supersonic combustor

This computational investigation provides a detailed comparative study of both non-reacting and reacting supersonic combustor flow-fields and attendant flow losses. Three different injection configurations in the same combustor are examined; a swept-sided ramp with base hydrogen injection, a straight ramp with base hydrogen injection and a thirty-degree downstream directed wall jet. Detailed comparisons are made with available reacting experimental data for the swept and the unswept ramps. A seven reaction, seven species reaction model as well as a global (one reaction) model are used for all three reacting cases. Relative performance (as measured by thrust) of all cases are described. Details of the flow, mixing, and combustion processes are discussed.

Riggins, David W.↗

NASA NTRS1990

Experimental and computational study of the effect of shocks on film cooling effectiveness in scramjet combustors

This paper presents results from a study conducted to investigate the effect of incident oblique shocks on the effectiveness of a coolant film at Mach numbers, typical of those expected in a scramjet combustor at Mach 15 to 20 flight. Computations with a parabolic code are in good agreement with the measured pressures and heat fluxes, after accounting for the influence of the shock upstream of its point of impingement on the plate, and the expansion from the trailing edge of the shock generator. The test data shows that, for the blowing rates tested, the film is rendered largely ineffective by the shock. Computations show that coolant blowing rates five to ten times those tested are required to protect against shock-induced heating. The implications of the results to scramjet combustor design are discussed.

Kamath, Pradeep S.↗

NASA NTRS1990

A computational design tool for scramjet combustor film cooling and fuel mixing predictions

A methodology for the analysis of a film cooled combustor of a hydrogen-fueled, supersonic combustion ramjet (scramjet) is presented. The SHIP3D parabolic flow code, based on an extension of the pressure-correction method of Patankar and Spalding for compressible flow is used to compute the flow field in the combustor. The effect of film cooling on the overall vehicle performance is then calculated using the one-dimensional, multistep cycle code, CPIPE. Results are presented for a conceptual combustor at Mach 16 flight conditions. It is seen that film cooling can have a significant effect on the engine performance.

Kamath, Pradeep S.↗

NASA NTRS1989

Hypersonic CFD applications for the National Aero-Space Plane

Design and analysis of the NASP depends heavily upon developing the critical technology areas that cover the entire engineering design of the vehicle. These areas include materials, structures, propulsion systems, propellants, integration of airframe and propulsion systems, controls, subsystems, and aerodynamics areas. Currently, verification of many of the classical engineering tools relies heavily on computational fluid dynamics. Advances are being made in the development of CFD codes to accomplish nose-to-tail analyses for hypersonic aircraft. Additional details involving the partial development, analysis, verification, and application of the CFL3D code and the SPARK combustor code are discussed. A nonequilibrium version of CFL3D that is presently being developed and tested is also described. Examples are given of portion calculations for research hypersonic aircraft geometries and comparisons with experiment data show good agreement.

Richardson, Pamela F.↗

NASA NTRS1989

Numerical simulation of flow through the Langley parametric scramjet engine

The numerical simulation of a three-dimensional turbulent, reacting flow through the entire Langley parametric scramjet engine has been obtained using a piecewise elliptic approach. The last section in the combustor has been analyzed using a parabolized Navier-Stokes code. The facility nozzle flow was analyzed as a first step. The outflow conditions from the nozzle were chosen as the inflow conditions of the scramjet inlet. The nozzle and the inlet simulation were accomplished by solving the three-dimensional Navier-Stokes equations with a perfect gas assumption. The inlet solution downstream of the scramjet throat was used to provide inflow conditions for the combustor region. The first two regions of the combustor were analyzed using the MacCormack's explicit scheme. However, the source terms in the species equations were solved implicitly. The finite rate chemistry was modeled using the two-step reaction model of Rogers and Chinitz. A complete reaction model was used in the PNS code to solve the last combustor region. The numerical solutions provide an insight of the flow details in a complete hydrogen-fueled scramjet engine module.

Srinivasan, Shivakumar↗