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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Supersonic combustion of hydrogen injected perpendicular to a ducted vitiated airstream

The reaction of hydrogen injected into a supersonic vitiated airstream from perpendicular injectors equally spaced over opposite walls of a two-dimensional duct is experimentally investigated. Information is obtained in the form of pitot pressure and gas sample surveys of the duct exit flow and static pressures along the duct walls for several injector arrangements differing in number, spacing, and size of injectors and operating at two levels of equivalence ratio. The amount of the injected hydrogen reacted is deduced from the static pressure data using a one-dimensional theory and is correlated with relative injection pressure and injector spacing-to-diameter ratio. These results are used with a mixing distribution correlation derived from nonreacting hydrogen-air results to predict static pressure distributions.

Rogers, R. C.↗
NASA NTRS1986

Numerical study of hydrogen-air supersonic combustion by using elliptic and parabolized equations

The two-dimensional Navier-Stokes and species continuity equations are used to investigate supersonic chemically reacting flow problems which are related to scramjet-engine configurations. A global two-step finite-rate chemistry model is employed to represent the hydrogen-air combustion in the flow. An algebraic turbulent model is adopted for turbulent flow calculations. The explicit unsplit MacCormack finite-difference algorithm is used to develop a computer program suitable for a vector processing computer. The computer program developed is then used to integrate the system of the governing equations in time until convergence is attained. The chemistry source terms in the species continuity equations are evaluated implicitly to alleviate stiffness associated with fast chemical reactions. The problems solved by the elliptic code are re-investigated by using a set of two-dimensional parabolized Navier-Stokes and species equations. A linearized fully-coupled fully-implicit finite difference algorithm is used to develop a second computer code which solves the governing equations by marching in spce rather than time, resulting in a considerable saving in computer resources. Results obtained by using the parabolized formulation are compared with the results obtained by using the fully-elliptic equations. The comparisons indicate fairly good agreement of the results of the two formulations.

Chitsomboon, T.↗
NASA NTRS2003

Calculation of Supersonic Combustion Using Implicit Schemes

One of the technology goals of NASA for advanced space transportation is to develop highly efficient propulsion systems to reduce the cost of payload for space missions. Developments of rockets for the second generation Reusable Launch Vehicle (RLV) in the past several years have been focused on low-cost versions of conventional engines. However, recent changes in the Integrated Space Transportation Program to build a crew transportation vehicle to extend the life of the Space Shuttle fleet might suggest that air-breathing rockets could reemerge as a possible propulsion system for the third generation RLV to replace the Space Shuttle after 2015. The weight of the oxygen tank exceeds thirty percent of the total weight of the Space Shuttle at launch while the payload is only one percent of the total weight. The air-breathing rocket propulsion systems, which consume oxygen in the air, offer clear advantages by making vehicles lighter and more efficient. Experience in the National Aerospace Plane Program in the late 1980s indicates that scramjet engines can achieve high specific impulse for low hypersonic vehicle speeds. Whether taking a form of Rocket Based Combined Cycle (RBCC) or Turbine Based Combined Cycle (TBCC), the scramjet is an essential mode of operation for air-breathing rockets. It is well known that fuel-air mixing and rapid combustion are of crucial importance for the success of scramjet engines since the spreading rate of the supersonic mixing layer decreases as the Mach number increases. A factored form of the Gauss-Seidel relaxation method has been widely used in hypersonic flow research since its first application to non-equilibrium flows. However, difficulties in stability and convergence have been encountered when there is strong interaction between fluid motion and chemical reaction, such as multiple fuel injection problems. The present paper reports the results from investigation of the effect of modifications to the original algorithm on the performance for multiple injectors.

Yoon, Seokkwan↗
NASA NTRS2008

Numerical Simulations of a Co-Axial Supersonic-Combusting Free-Jet Experiment

CFD calculations using the Reynolds-averaged Navier-Stokes equations coupled with species continuity equations have been made for a supersonic coaxial-jet CFD-validation experiment to determine the sensitivity of the external flowfield to the main-nozzle exit profile. Four different nozzle exit profiles were used in the study: a uniform profile, one computed using only the nozzle geometry, one computed using the nozzle geometry and part of the upstream facility combustor, and one using the nozzle and the full facility combustor. Two cases were examined using the four profiles: a non-reacting case without coflow and a reacting case with hydrogen coflow. Results show that the nozzle exit profile has a significant effect on the external flowfield. The uniform profile produced the longest jet while the profile created with the full combustor produced the shortest jet. The nozzle-only and part-combustor profiles fell between the other two profiles. The reacting flow was found to be more sensitive to the nozzle exit profile since it affects the downstream mixing and combustion. These calculations indicate the importance of properly setting the nozzle-exit profile for this type of calculation.

Gaffney, Richard L.↗
NASA NTRS2001

Performance of Several High Order Numerical Methods for Supersonic Combustion

The performance of two recently developed numerical methods by Yee et al. and Sjoegreen and Yee using postprocessing nonlinear filters is examined for a 2-D multiscale viscous supersonic react-live flow. These nonlinear filters can improve nonlinear instabilities and at the same time can capture shock/shear waves accurately. They do not, belong to the class of TVD, ENO or WENO schemes. Nevertheless, they combine stable behavior at discontinuities and detonation without smearing the smooth parts of the flow field. For the present study, we employ a fourth-order Runge-Kutta in time and a sixth-order non-dissipative spatial base scheme for the convection and viscous terms. We denote the resulting nonlinear filter schemes ACM466-RK4 and WAV66-RK4.

Sjoegreen, Bjoern↗
NASA NTRS1986

Numerical study of supersonic combustion using a finite rate chemistry model

The governing equations of two-dimensional chemically reacting flows are presented together with a global two-step chemistry model for H2-air combustion. The explicit unsplit MacCormack finite difference algorithm is used to advance the discrete system of the governing equations in time until convergence is attained. The source terms in the species equations are evaluated implicitly to alleviate stiffness associated with fast reactions. With implicit source terms, the species equations give rise to a block-diagonal system which can be solved very efficiently on vector-processing computers. A supersonic reacting flow in an inlet-combustor configuration is calculated for the case where H2 is injected into the flow from the side walls and the strut. Results of the calculation are compared against the results obtained by using a complete reaction model.

Chitsomboon, T.↗
NASA NTRS1973

Analytical and experimental study of supersonic combustion of hydrogen in a vitiated airstream

Detailed probe measurements of total temperature, pressure, and composition were taken in a two-dimensional test section 35.6 cm downstream of hydrogen injection. A high pressure gas generator supplied Mach 2.44 vitiated air or inert gas at elevated temperatures and at a static pressure equal to that of the hydrogen. Special water-cooled probes and sampling techniques were developed for the short test times required by heat-sink hardware. Independent methods of measuring stream total temperatures are compared. For the pure mixing case, the computed composition profile agreed well with the experimental profile. The analysis takes into account the wall boundary layer and the initial boundary layer in the main stream. Ignition of hydrogen, as determined from photographic exposures of the radiating gases, varied from 30 to 10 cm downstream from injection for a 45 K increase in local free-stream static temperature.

Burrows, M. C.↗
NASA NTRS1985

Supersonic combustion of a silane/methane mixture

A test program was conducted to evaluate the ignition and combustion characteristics of a 20 volume percent silane/methane mixture using a model scramjet combustor with Mach 2 combustor entrance conditions. The test gas total temperature was varied from 1200 to 3900 R. The mixture autoignited at a total temperature of 1650 R. This autoignition temperature can be contrasted with 2330 R for hydrogen and 1350 R for a 20 percent silane/hydrogen mixture in similar hardware. Methane without the silane additive did not autoignite in this configuration at temperatures as high as 3900 R.

Northam, G. B.↗
NASA NTRS1987

Design and evaluation of a new injector configuration for supersonic combustion

Ignition and flameholding behavior data are presented from experiments with a novel scramjet injector configuration consisting of five small upstream pilot fuel injectors, a rearward-facing step, and three primary fuel injectors downstream of the step; all injection is perpendicular to the Mach 2 airflow, and the ignition source is located in the recirculation region downstream of the step. The design attempts to minimize the amount of igniter energy required for ignition, but without introduction of the large losses associated with bluff bodies. Argon plasma, argon-hydrogen plasma, pyrophoric silane/hydrogen, and surface-discharge ignition sources are tested. The injector is found to be able to hold flames at 1400 R total temperatures without the use of large bluff bodies; the argon-hydrogen plasma was the most effective igniter.

Wagner, Timothy C.↗
NASA NTRS1989

Initial supersonic combustion facility measurements

A combustion test tunnel designed for continuous operation to 2000 K was assembled. Flow quality of a Mach 2 nozzle for use with this tunnel was examined using an array of impact probes. The performance of gas shields used to protect optical windows was examined using both shadowgraphs and planar laser induced iodine fluorescence. High speed videography was used to aid in design of pressure relief panels related to hydrogen combustion testing safety.

Krauss, Roland H.↗
NASA NTRS1989

Evaluation of parallel injector configurations for supersonic combustion

Techniques for the enhancement of mixing in a scramjet combustor are presently investigated experimentally with wall-mounted parallel injector ramps that are speculated to prove useful at high speeds in the extraction of thrust from engine- or airframe-coolant hydrogen. The injector ramps were designed to generate a reflected shock wave from the duct top wall which will pass just downstream of the fuel injectors' barrel shock; fuel injector Mach number is 1.7, in virtue of which an underexpanded fuel flow is generated. Shadowgraph and UV-TV flow visualizations were conducted for three duct configurations.

Northam, G. Burton↗
NASA NTRS1991

Navier-Stokes simulation of the supersonic combustion flowfield in a ram accelerator

A computational study of the ram accelerator, a ramjet-in-tube device for accelerating projectiles to ultrahigh velocities, is presented. The analysis is performed using a fully implicit TVD scheme that efficiently solves the Reynolds-averaged Navier-Stokes equations and the species continuity equations associated with a finite rate combustion model. Previous analyses of this concept were based on inviscid assumptions. The present results indicate that viscous effects are of primary importance; in all the cases studied, shock-induced combustion always started in the boundary layer. The effects of Mach number, mixture composition, pressure, and turbulence are investigated for various configurations. Two types of combustion processes, one stable and the other unstable, were observed depending on the inflow conditions. In the unstable case, a detonation wave is formed, which propagates upstream and unstarts the ram accelerator. In the stable case, a solution that converges to steady-state is obtained, in which the combustion wave remains stationary with respect to the ram accelerator projectile. The possibility of stabilizing the detonation wave by means of a backward facing step is also investigated. In addition to these studies, two numerical techniques were tested. These two techniques are vector extrapolation to accelerate convergence, and a diagonal formulation that eliminates the expense of inverting large block matrices that arise in chemically reacting flows.

Yungster, Shaye↗