Showing papers on "Scramjet published in 1994"

Advanced Injection and Mixing Techniques for Scramjet Combustors

Abstract: Scramjet combustor fuel injection and mixing enhancement techniques are reviewed. The injection techniques include hole injection from the combustor wall, slot injection parallel to the flow, and injection from struts and the rear of ramps. Three new advanced mixing techniques are presented. The first is a combustor, curved so that buoyancy forces will aid in the penetration of the fuel across the combustor. The second is pulsation of the fuel injectors to increase penetration and mixing. A fluidic technique, a modified Hartmann-Sprenger tube, is identified as a strong candidate to generate the pulsations. The third is the injection behind pylons to allow deep penetration into the air stream. This technique is likely to produce high base pressures on the injector structure, particularly if base burning is encouraged. Curved or slanted pylons can be used to increase the recovery of fuel jet momentum. The potential of the new mixing techniques to increase scramjet engine performance is assessed.

Supersonic Combustion of Hydrogen/Air in a Scramjet Combustion Chamber.

Abstract: Due to the importance of supersonic combustion in hypersonic propulsion an experimental investigation of the flowfield of a supersonic ramjet combustion chamber (scramjet) at a Mach number of 2 has been carried out. The simulated free flight Mach number was about 5. The hydrogen was injected parallel to the air stream throughthe base of a symmetric wedge (strut) centered in the two dimensionalcombustion chamber. The strut served as a flameholder. Spontaneous OH-emission, schlieren photography and pitot pressure measurements were used to acquire basic information about the flowfield. Quantitativ measurements were made with a 2D Laser-Doppler- Velocimeter (LDV) from which the mean velocity, turbulent intensitiesand Reynolds stresses were determined. Coherent-Anti-Stokes-Raman- Scattering (CARS) was used to determine the temperature histograms for each measurement point. The histrograms show the turbulent temperature fluctuations in the combustion chamber.

Effect of film cooling/regenerative cooling on scramjet engine performances

Abstract: Film cooling was modeled to allow performance prediction of scramjet engine design. The model was based on experimental results of the compressible mixing layer for the vicinity of the injection slot, and on analytical results of the turbulent boundary layer in the region far from the slot. The film cooling model was integrated with a quasi-one-dimensional scramjet performance prediction model. In an engine employing a combination of film cooling and regenerative cooling, coolant flow rate of the engine slightly exceeded the stoichiometric flow rate, even at high flight Mach numbers, and had the best specific impulse and system pressure performances. These advantages were achieved by an increase in the volume flow rate and a decrease in the velocity difference between the main flow and the coolant, both due to an increase in the film coolant temperature. The effective cooling system with a combination of film cooling and regenerative cooling was also advantageous with regard to avoidance of excess cooling of the engine wall.

Review of shock-induced supersonic combustion research and hypersonic applications

Abstract: This article focuses on research in supersonic combustion and combustion kinetics in high-speed flow between 1959-1968, and the application of the experimental results to hypersonic propulsion. The analysis discusses both advantages and problems for premixing the fuel and employing shock-induced combustion as an ignition method for a scramjet flying at a high Mach number. The experimental tests are discussed, including implications to the chemical kinetics of the high-velocity combustion process. The conditions were confined to relatively low pressure, less than 2 atm (200 kPa). The results were considered to be mainly applicable for high-altitude scramjet flight, at low static pressure, where chemical reaction distances will be long. At these lower pressures, "shock-induced combustion" may be the predominant effect in a scramjet application, and it has some advantages that are discussed. The relation between shock-induced combustion and "detonation" is also discussed. In addition, an attempt is made to resolve the conflicting experimental data published in the 1960s relating to "standing detonation waves" and shock-induced combustion.

Suppression and Enhancement of Mixing in High-Speed Reacting Flow Fields

Abstract: Work is underway at the NASA Langley Research Center to develop a hydrogen-fueled supersonic combustion ramjet, or scramjet, that is capable of propelling a vehicle at hypersonic speeds in the atmosphere. Recent research has been directed toward the optimization of the scramjet combustor and, in particular, the efficiency of fuel-air mixing and reaction taking place in the engine. With increasing Mach number, the degree of fuel-air mixing through natural convective and diffusive processes is significantly reduced leading to an overall decrease in combustion efficiency and thrust. Even though the combustor flow field is quite complex, it can be viewed as a collection of spatially developing and reacting supersonic mixing layers or jets from fuel injectors mixing with air, one of which serves as an excellent physical model for the overall flow field. This work is focused on understanding the mechanisms of mixing (or lack thereof) and on the development of techniques for its enhancement in compressible turbulent reacting flows. Results generated by direct numerical simulations (DNS) are first used to demonstrate the mechanisms for reduced mixing in shear layers. To counter the effects of suppressed mixing, several mixing enhancement techniques are then discussed. The most successful approaches involve longitudinal vorticity induced into the flow field. Several means for inducing vorticity are studied and assessed.

Wind-tunnel blockage and actuation systems test of a two-dimensional scramjet inlet unstart model at Mach 6

Abstract: The present study examines the wind-tunnel blockage and actuation systems effectiveness in starting and forcibly unstarting a two-dimensional scramjet inlet in the NASA Langley 20-Inch Mach 6 Tunnel The intent of the overall test program is to study (both experimentally and computationally) the dynamics of the inlet unstart; however, prior to the design and fabrication of an expensive, instrumented wind-tunnel model, it was deemed necessary first to examine potential wind-tunnel blockage issues related to model sizing and to examine the adequacy of the actuation systems in accomplishing the start and unstart The model is equipped with both a moveable cowl and aft plug Windows in the inlet sidewalls allow limited optical access to the internal shock structure; schlieren video was used to identify inlet start and unstart A chronology of each actuation sequence is provided in tabular form along with still frames from the schlieren video A pitot probe monitored the freestream conditions throughout the start/unstart process to determine if there was a blockage effect due to the model start or unstart Because the purpose of this report is to make the phase I (blockage and actuation systems) data rapidly available to the community, the data is presented largely without analysis of the internal shock interactions or the unstart process This series of tests indicated that the model was appropriately sized for this facility and identified operability limits required first to allow the inlet to start and second to force the unstart

Technology and staging effects on two-stage-to-orbit systems

Abstract: Horizontal takeoff and landing two-stage systems with an airbreathing first stage and rocket second stage are evaluated for staging Mach numbers that range from 5 to 14. All systems are evaluated with advanced technologies being developed in the NASP Program and sized to the same mission requirements. With these advanced technologies, the two-stage systems are heavier than the single stage. The weights of the two-stage systems are closely related to staging. Using a rocket on the first stage to accelerate from the turboramjet limit of Mach 6 to Mach 10 signiificantly decreases dry weight as compared to the Mach 6-staged system. The optimum dry weight staging Mach number for the scramjet two-stage system is Mach 12. At a 40 percent weight growth (current technology level), the scramjet two-stage systems are half the weight and less sensitive to weight changes than the single stage, but still require substantial technology development in the areas of inlets, nozzles, ramjet propulsion, active cooling, and high-temperature structures.

Modeling Scramjet Combustor Flowfields with a Grid Adaptation Scheme

Abstract: 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.

Thrust loss due to supersonic mixing

Abstract: The maximum thrust developed by a device in which two streams mix in a parallel configuration at supersonic velocities is estimated. Total pressure profiles in a two-dimensional, compressible shear layer are calculated by assuming turbulent Prandtl and Lewis numbers of unity. As the convective Mach number Mc rises, the total pressure acquires a defect that becomes large for Mc > 1. For shear layers with equal freestream total pressures, an analytical relation for the defect vs Mc is found. The extent and magnitude of the defect agrees well with experimental data. The loss in total pressure is connected to the loss in thrust of a simplified model of a scramjet. The thrust loss is about 30% for Mc = 2 and 50% for Mc 3. The trends are insensitive to details of the shear-layer velocity profile and to the ratios of freestream quantities. The role of turbulent-energy dissipation in the reduction of total pressure is discussed.

Three-dimensional analysis of scramjet nozzle flows

Abstract: Numerical and experimental results of performance in National Aero-Space Plane-like nozzles are compared. The full Navier-Stokes equation with the Baldwin-Lomax turbulence model was adopted to solve the threedimensional nozzle flows. A code validation study was conducted using pressure and heat-flux distributions measured. The interactions of separation shocks with the main internal flow under overexpanded conditions were investigated. The interaction yields higher performance in scramjet nozzles than that estimated assuming a two-dimensional separation. The losses in the nozzle internal flow and the overexpansion loss were evaluated.

Air-Breathing Hypersonic Cruise: Prospects for Mach 4–7 Waverider Aircraft

Abstract: In the Mach 4-7 range, waverider aircraft are considered as candidates for both short- and long-range cruise missions, as hypersonic missiles, and as high L/D highly maneuverable craft. The potential for near- and far-term application of airbreathing engines to the waverider vehicle missions and concepts is presented. Attention is focused on the cruise mission and attempts are made to compare and contrast it with the accelerator mission.

Experimental and computational results for the external flowfield of a scramjet inlet

Abstract: A two-dimensional hypersonic scramjet inlet has been investigated in a combined experimental and analytical program aimed at addressing the fundamental issues related to the design of scramjet inlets. The experimental portion of the program was conducted in the Calspan 48-in. shock tunnel at Mach 10 and 13. The computational analysis was conducted using a two-dimensional parabolized Navier-Stokes computational fluid dynamics (CFD) code. This article addresses the issues concerned with the flow over the external forebody of the inlet which consists of a blunted wedge followed by an isentropic compression. The pressure and heat transfer distributions over the forebody are investigated for ranges of Reynolds number, Mach number, wall-to-freestream temperature ratio, and nose bluntness. Comparison of the test results and CFD predictions show that good agreement for the heat transfer distributions is achieved. However, the predicted pressure distribution on the forward blunted wedge was consistently underpredicted by 18-33% relative to the experimental measurements. Several phenomena were investigated in an attempt to explain the discrepancy between predicted and measured pressure distributions, including classical viscous leading-edge interactions, blunt leading-edge interactions, slip flow effects, flow condensation, flow angularity, and facility Mach number uncertainty. Although the discrepancy in the forebody pressure ratios could be caused by a combination of the factors listed above, a deficiency in the modeling of the viscous interaction region by the CFD codes and facility flow angularity are shown to be the strongest contributors.

Experimental studies on mixing of two co-axial high-speed streams

Abstract: Currently, the focus of attention in aerospace propulsion research has been on the development of advanced air-breathing propulsion systems like scramjets, air-augmented rockets, multimode engines for the hyperplane, etc. A crucial technical problem to be tackled in such systems is the enhancement of mixing between two highspeed gaseous streams. Various methods have been tried so far, but the field is still very much an open one. Tests conducted earlier on subsonic flow mixing have shown that large-scale secondary flows, and not viscous diffusion, are the key to quick, low-loss, efficient mixing. In this regard, a lobe-type, convergent-divergent, supersonic nozzle, named the "Petal" nozzle, was designed, fabricated, and tested. Near-complete mixing with low total momentum loss within a short mixing chamber of length to diameter ratio of 4.35 was achieved.

Numerical Study of the Performance of Swept, Curved Compression Surface Scramjet Inlets

Abstract: A computational performance enhancement study was performed employing systematic modifications to a planarsidewall compression scramjet inlet operating at an entrance Mach number of 4 and at a dynamic pressure of 2040 psf. The variations included modifying the planar-side wall compression angle as a function of height, utilizing sidewall curvature, and employing, simultaneously, both forward-swept and reverse-swept compression surfaces. Turbulent flowfield solutions were generated by solving the Reynolds-averaged Navier-Stokes equations to obtain inlet performance parameters such as total-pressure recovery, mass capture, and flowfield pressure distortion (the ratio of maximum static pressure to minimum static pressure generated at the inlet exit plane). Additionally, an inviscid parametric study was performed by employing solutions to the Euler equations to optimize a cubic polynomial that defined the longitudinal sidewall geometry. A final viscous flowfield solution of the optimized inviscid inlet geometry yielded inlet performance improvements; however, inlet top-wall surface boundary-layer shock wave separation interactions persisted. Hence, this numerical study demonstrated that enhanced performance is obtainable via curved-wall geometric modifications to the standard planar-sidewall inlet design, although future work should employ constraints to mitigate detrimental flow separation effects.

Integrated inlets for cone-derived waveriders with combined transverse and longitudinal curvature

Abstract: A generic aerospace vehicle is constituted of three parts: a forebody, a scramjet, and an afterbody. This paper is concerned with the design of the forebody and the inlet by means of waverider configurations. Hypersonic stream surfaces past a cone with combined transverse and longitudinal curvature are used to design integrated inlets for wavenders. By suitably choosing a second-order even polynomial stream surface, inlet shapes can be controlled. Mass flow rate, wetted surface area, volume, lift, drag, and liftto-drag ratio can be found in closed forms. Effects of various parameters on the shape of the inlet are found and discussed in detail. Thus, an overall aerodynamic design of a hypersonic vehicle can be established in a simple and systematic way. Flight conditions are determined for a wide range of shapes and parameters as well as other related factors depending on the designer's interest.

Scramjet fuel injection system having independent fuel supplies for supersonic and hypersonic operation

Abstract: A fuel injector for a scramjet engine includes independently manifolded normal (48, 54) and axial fuel injection (56). Various construction details are developed that provide effective performance of the fuel injector over a range of speed and combustor conditions. In a particular embodiment the fuel injector has a wedge shape with a swept leading edge. The fuel injector includes a pair of sidewalls, a top surface, a downstream facing surface, and a plurality of apertures in the sidewalls, top surface and downstream facing surface. At speeds less than hypersonic (normal injection), fuel is injected into the combustor from the apertures in the sidewalls (48) and top surface (54). At hypersonic speeds (axial injection), fuel is injected through the apertures in the downstream facing surface (56).

Shock induced mixing and combustion in scramjets

Abstract: Analytical and experimental investigations of shock induced mixing and combustion have been conducted in the present study to determine its usefulness in enhancing mixing and combustion in scramjets. Analytical techniques were developed to describe the steady interaction of an oblique shock wave and a planar mixing region. Under the assumed conditions, it was ascertained that the trajectory of a shock wave through a mixing region is a function of the Mach number distribution. Typically, when an oblique shock wave enters a region of lower Mach number, the shock and flow deflection angles increase, and the post-shock pressure decreases. Due to the importance of vorticity in mixing problems, an analytical expression for the vorticity jump across the shock wave was also obtained. The analysis shows that to be assured of an amplification of the pre-shock vorticity, the gradients of velocity and density within the mixing region must lie in the same direction. These results have important practical implications for the optimisation of shock induced mixing in scramjets. To test the analytical techniques and examine shock induced mixing and combustion, a number of experiments were conducted in The University of Queensland T4 shock tunnel facility (at Mach numbers between approximately 5.0 and 6.6), and the University of Oxford gun tunnel (at a Mach number of 7. 1). Planar ducts with central strut injection and a variety of shock inducing wedge angles (0, 5, 10, and 15°) were used in both facilities, and additional shock tunnel experiments were conducted using a complex engineering scramjet configuration. It was determined that the shock wave-mixing region analysis provided a good prediction of the experimentally observed interaction process. Turbulent fluctuations within the mixing regions caused the shock waves to break up into a number of unsteady paths which persisted into the free stream on the other side of the layers. Using the vorticity analysis, mixing enhancement was predicted since the calculated vorticity amplification was between approximately 100 and 400 % fo r the different shock wave-mixing region interactions. Mixing augmentation through shock wave impingement was observed immediately after shock processing in both the gun tunnel and shock tunnel --, experiments. Shock induced combustion was observed in the complex engineering scram jet model, and in the 1 0° wedge and duct configuration. From the current research, it is apparent that shock induced mixing and combustion remain attractive possibilities for the enhancement of mixing and combustion in scramjets.

"Investigation of the Flowfield of a Scramjet Combustor with Parallel H2-Injection through a Strut by Particle Image Displacement Velocimetry."

Abstract: The flow field in a supersonic combustor with H2-injection through a strut is investigated by particle image displacement velocimetry. The flowfield is characterized by shock trains which originate at the leading edge of the strut, an expansion originating at the interaction of the fuel and the air flows. PIV recordings are taken without and with ignition of the fuel/oxidizer mixture. Velocities up to 900 m/s were measured. The measured velocity fields are compared with the flow structures in the shadow photographs.

Investigation of ramp injectors for supersonic mixing enhancement

Abstract: A comparative study of wall mounted swept ramp injectors fitted with injector nozzles of different shape has been conducted in a constant area duct to explore mixing enhancement techniques for scramjet combustors. Six different injector nozzle inserts, all having equal exit and throat areas, were tested to explore the interaction between the preconditioned fuel jet and the vortical flowfield produced by the ramp: circular nozzle (baseline), nozzle with three downstream facing steps, nozzle with four vortex generators, elliptical nozzle, tapered-slot nozzle, and trapezoidal nozzle. The main flow was air at Mach 2, and the fuel was simulated by air injected at Mach 1.63 or by helium injected at Mach 1.7. Pressure and temperature surveys, combined with Mie and Rayleigh scattering visualization, were used to investigate the flow field. The experiments were compared with three dimensional Navier-Stokes computations. The results indicate that the mixing process is dominated by the streamwise vorticity generated by the ramp, the injectors' inner geometry having a minor effect. It was also found that the injectant/air mixing in the far-field is nearly independent of the injector geometry, molecular weight of the injectant, and the initial convective Mach number.

Measurements of scramjet thrust in shock tunnels

Abstract: By using results obtained in tests on supersonic combustion of hydrogen in air, the conditions governing model size and operating pressure levels for shock tunnel experiments on models of flight vehicles with scramjet propulsion are established. It is seen that large models are required. The development of the stress wave force balance is then described, and its use as a method of measuring thrust/drag on such models is discussed. Test results on a simple, fully integrated scramjet model, with intakes, combustion chambers, thrust surfaces and exterior surfaces, using a 13 percent silane 87 percent hydrogen fuel mixture, showed that a steady state with thrust generation could be achieved within the shock tunnel test time, and the thrust could be measured. Results are presented for a range of stagnation enthalpies, and show that the scramjet model produces net positive thrust at velocities up to 2.4 km/sec.

Calculation of scramjet inlet with thick boundary-layer ingestion

Abstract: Numerical flowfields around a scramjet inlet model are simulated and analyzed. The present inlet flowfield is characterized by thick boundary-layer ingestion and strong viscous/inviscid interaction because of a combined effect of high hypersonic freestream Mach and low Reynolds numbers. Shock-induced separation further enlarges regions of viscous flows which occupy most of the inlet flowfield. Results obtained from the computations with the PARC code developed for ideal gas are presented for several 2D cases at various hypersonic Mach numbers ranging from 10 to 25, and two 3D simulations at Mach numbers of 12 and 19 are also discussed. Comparison between computation and experiment is made in terms of pressure distributions at the wall center line. Large discrepancy is observed and may be partially attributed to the lack of real gas and/or 3D effects in the simulation as well as to the uncertainty of the experiment.

Hypersonic Combustion — Status and Directions

Abstract: Man’s imagination has long been sparked by flight. Today sustained manned flight at hypersonic speed remains the largest unexplored region of the possible flight envelope. True aerospace planes flying at near orbital speeds will require a new form of airbreathing engine — the supersonic combustion ramjet or scramjet — and will also be required to “break the thermal barrier” with active fuel cooling of the engine and vehicle to survive. This paper will briefly review the internal flow environment of scramjets and discuss some of the relevant issues which arise in attempts to produce meaningful ground simulations of these flows. Some recent computational studies of the potential influence of injector design variables on performance will be reviewed as an example of the application of CFD to hypersonic combustion modeling. In addition, some projections to the future as to where emphasis should be placed in developing experimental capability and in extending computer modeling will be put forward.