Scramjet Test Facility Research Articles

Non-Reacting Examination on Transverse Injection Upstream of Aft Wall Angled Cavities in an Axisymmetric Supersonic Flow

Experiments on an air-fueled non-reacting scramjet test facility with four distinct aft wall angled cavity designs were performed in a Mach 1.8 flow field to explore the effects of transverse fuel injection system on scramjet combustors. Axisymmetric combustor cavity models were featured with two consecutive angles being inclined towards the downstream flow direction at the rear end. The wall mounted injector was located upstream from the cavity at a distance of 10 mm. Mixing performance within the various rear wall angled cavities were evaluated and compared with No-injection case for different injection pressures. Results revealed that the flow within the combustor was more uniform mixing compared to No-cavity configuration and was greatly influenced by injection pressures. Increase in injection pressures were characterized by enhanced mixing, greater stagnation pressure loss values and unstable flow.

Dual-Pump Coherent Anti-Stokes Raman Spectroscopy Measurements in a Dual-Mode Scramjet

In this paper, the authors describe dual-pump coherent anti-Stokes Raman spectroscopy (CARS) measurements of mixing and combustion in a direct-connect scramjet combustor operating at equivalent flight Mach typical of the ramjet–scramjet transition, in the scram mode. Measurements were performed in the University of Virginia’s scramjet test facility in which the air is heated by electrical resistance heaters. The CARS technique is used to acquire temporally and spatially resolved measurements of temperature and species mole fraction. Measurements were at four planes: one upstream of an fuel injector and three downstream. Contour plots of mean flow and standard deviation statistics are presented for cases with and without reaction of the fuel. The vibrational temperature at the exit of the facility nozzle, and in the freestream of the scramjet combustor, is elevated compared with the rotational temperature; the air– vibrational temperature is the same as the facility stagnation temperature. There are spatial nonuniformities of temperature exiting the heater. The mixing of the fuel jet from the single ramp wall injector and the growth of the plume downstream is shown. The flame is attached at the injector and propagates from the wall adjacent to the fuel plume, around the periphery of the plume, before engulfing the whole plume. Computational fluid dynamics modeling shows that the flow can be predicted well using hybrid large-eddy simulation/Reynolds-averaged Navier–Stokes methods with relatively simple subgrid models, provided facility effects are properly accounted for.

Experimental Study on Effects of Fuel Injection on Scramjet Combustor Performance

In order to investigate the effects of fuel injection distribution on the scramjet combustor performance, there are conducted three sets of test on a hydrocarbon fueled direct-connect scramjet test facility. The results of Test A, whose fuel injection is carried out with injectors located on the top-wall and the bottom-wall, show that the fuel injection with an appropriate close-front and centralized distribution would be of much help to optimize combustor performances. The results of Test B, whose fuel injection is performed at the optimal injection locations found in Test A, with a given equivalence ratio and different injection proportions for each injector, show that this injection mode is of little benefit to improve combustor performances. The results of Test C with a circumferential fuel injection distribution displaies the possibility of ameliorating combustor performance. By analyzing the effects of injection location parameters on combustor performances on the base of the data of Test C, it is clear that the injector location has strong coupled influences on combustor performances. In addition, an inner-force synthesis specific impulse is used to reduce the errors caused by the disturbance of fuel supply and working state of air heater while assessing combustor performances.

Numerical simulation of model scramjet combustor flowfield

AbstractA new concept has been raised and adopted in this paper to enlarge the scope of the two‐dimensional model particularly for the purpose of dealing with three‐dimensional normal injection cases. Meanwhile, the method has a very good performance for its short cyclic period. The new idea was realized through special resolution with continuity equations; i.e., mass flow was directly added in the source term of the continuity equation. To prove the robustness of this illuminating method, comparisons using calculations were carried out, and the results are satisfactory. A model scramjet combustor tested on the free‐jet scramjet test facility was illustrated and underwent numerical calculations with the two‐dimensional program, adopting the above simplified injecting method. To simulate the chemical reaction process in the scramjet tunnel, a five‐species, single‐step reaction model was introduced in the calculation process. This research presents the major aerodynamic parameters and components of mass fraction distribution within the model combustor channel, which made it easy to observe and analyze the flowfield. Finally, wall pressure comparisons between the numerical and experimental results were carried out to verify the accuracy of the calculation model. © 2007 Wiley Periodicals, Inc. Heat Trans Asian Res, 36(5): 295– 302, 2007; Published online in Wiley InterScience (www.interscience.wiley.com). DOI 10.1002/htj.20159

Design, Tests, and Verification of a Diffuser System for a Mach 4 Scramjet Test Facility

R directed toward developing the technology for an airframe-integrated modular scramjet engine concept currently being conducted at NASA Langley Research Center, involves experimental investigations of hydrogen-burning scramjet models at simulated Mach 4 and 7 flight conditions. The present concept of the airframeintegrated scramjet engine has modules of rectangular cross section with swept leading edges that produce an asymmetric downward flow that sharply increases locally near the cowl when the scramjet inlet unstarts. Tests on a heat-sink, hydrogen-burning model representing one module of this concept have been conducted at Langley in a facility which duplicates Mach 7 flight conditions. For Mach 4 tests, an engine test cell is being modified to contain a freejet, blowdown tunnel that will exhaust to the atmosphere. To provide shock-free tunnel flow to the modular scramjet engine and to meet the atmospheric exhaust condition, a diffuser system is required. The available mass flow rates and the size of existing tunnel hardware dictated that the scramjet model block up to 33% of the tunnel nozzle exit area. A diffuser preliminary design was defined based on published experimental results. However, most of the literature results are not directly applicable to conditions where the model blockage is as high as 33%, and where the model design produces asymmetry by creating a downstream flow that increases when the model inlet unstarts. An experimental investigation was, therefore, undertaken using unheated air and a subscale model of the tunnel-scramjetdiffuser system. A description of the subscale system and some data results obtained during tests with this system are presented. These results were subsequently verified, as a data comparison will show, by full-scale engine combustion tests in a Mach 4 facility utilizing a diffuser system based upon the best arrangement of the subscale tunnel scramjetdiffuser system.