Showing papers on "Scramjet published in 2006"

Flight Data Analysis of the HyShot 2 Scramjet Flight Experiment

Abstract: The development of scramjet propulsion for alternative launch and payload delivery capabilities has been composed largely of ground experiments for the last 40 years. With the goal of validating the use of short duration ground test facilities, a ballistic reentry vehicle experiment called HyShot was devised to achieve supersonic combustion in flight above Mach 7.5. It consisted of a double wedge intake and two back-to-back constant area combustors; one supplied with hydrogen fuel at an equivalence ratio of 0.34 and the other unfueled. Of the two flights conducted, HyShot 1 failed to reach the desired altitude due to booster failure, whereas HyShot 2 successfully accomplished both the desired trajectory and satisfactory scramjet operation. Postflight data analysis of HyShot 2 confirmed the presence of supersonic combustion during the approximately 3 s test window at altitudes between 35 and 29 km. Reasonable correlation between flight and some preflight shock tunnel tests was observed.

Characteristics of Dual Transverse Injection in Scramjet Combustor, Part 1: Mixing

Abstract: The mixing characteristics of a dual transverse injection system in a scramjet combustor were studied with numerical methods. The effects of the jet-to-cross-flow momentum flux ratio and the distance between injectors on mixing characteristics were investigated. Three-dimensional Navier–Stokes equations, including the k–! SST turbulencemodel,weresolvedwiththe finitevolumemethodadoptingtheupwindmethodofEdwards’lowdiffusion flux splitting scheme. It is shown that the mixing characteristics of a dual transverse injection system are very differentfromthoseofasingleinjectionsystem.Therearinjection flowisstronglyinfluencedbyblockageeffectsdue to the momentum flux of the front injection flow and thus has higher expansion and penetration than the front injection flow. The dual injection system has a higher mixing rate and a higher penetration but have more losses of stagnation pressure than the single injection system. It is also shown that there is an optimal distance between injectors for mixing characteristics and that the optimal distance increased as the jet-to-cross-flow momentum flux ratio increased.

Propulsive performance of airbreathing pulse detonation engines

Abstract: The propulsive performance of airbreathing pulse detonation engines at selected flight conditions is evaluated by means of a combined analytical/numerical analysis. The work treats the conservation equations in axisymmetric coordinates and takes into account finite-rate chemistry and variable thermophysical properties for a stoichiometric hydrogen/air mixture. In addition, an analytical model accounting for the state changes of the working fluid in pulse detonation engine operation is established to predict the engine performance in an idealized situation. The system under consideration includes a supersonic inlet, an air manifold, a valve, a detonation tube, and a convergent-divergent nozzle. Both internal and external modes of valve operation are implemented. Detailed flow evolution is explored, and various performance loss mechanisms are identified and quantified. The influences of all known effects (such as valve operation timing, filling fraction of reactants, nozzle configuration, and flight condition) on the engine propulsive performance are investigated systematically. A performance map is established over the flight Mach number of 1.2-3.5. Results indicate that the pulse detonation engine outperforms ramjet engines for all the flight conditions considered herein. The benefits of pulse detonation engines are significant at low-supersonic conditions, but gradually decrease with increasing flight Mach number.

Experimental and numerical study of a dual-mode scramjet combustor

Abstract: A Mach 2, hydrogen-air combustor with an unswept 10-deg ramp fuel injector was experimentally and numerically studied for a simulated flight Mach number near 5. Numerical modeling was performed using the Viscous Upwind Algorithm for Complex Flow Analysis code, and results were compared against experimental wall-pressure distributions, fuel plume images, and fuel plume velocity measurements. The model matched wall-pressure distributions well for the case of fuel-off and fuel-air mixing. For a fuel-air reacting case, pressure was matched well in the upstream third of the duct. Downstream, however, the pressure rise as a result of combustion was underpredicted. Based on the fuel plume imaging and velocity measurements, fuel plume shape was matched well for both the mixing and reacting cases. However, plume size, penetration, and centerplane axial growth were generally underpredicted by the model. The full extent of the velocity reduction caused by thermal choking was also not predicted. Despite these findings, the numerical model performed better than a previous model developed by the investigators. It was proposed that differences between the present numerical model and experiment stemmed from numerical underprediction of fuel-air turbulent mixing, and this resulted in underprediction of heat release.

Characteristics of Dual Transverse Injection in Scramjet Combustor, Part 2: Combustion

Abstract: The combustion characteristics of a dual transverse injection system in a scramjet combustor were studied with numerical methods. The effects of the jet-to-crossflow momentum flux ratio and the distance between injectors on combustion characteristics were investigated. It is shown that the dual injection system has very different combustion characteristics with respect to the single injection system; the burning process of the rear injection flow is strongly influenced not only by the blockage effects but also by the preheating effects due to the chemical reactions of the front injection flow. The dual injection system has a higher burning rate and a higher flame height but more loss of stagnation pressure than the single injection system. It is also shown that there is an optimal distance between injectors for combustion characteristics and that the optimal distance increases as the jet-to-crossflow momentum flux ratio increases.

Performance of a Dual-Mode Combustor with Multistaged Fuel Injection

Abstract: A scramjet combustor, which has a constant cross-sectional area combustor (designated as combustor 1), and a diverging combustor (designated as combustor 2), was examined with both single and two-stage fuel injection to determine whether it could be operated in the ramjet mode. The combustor was directly connected to a facility nozzle, the Mach number at the exit of the nozzle being 2.4. Total temperature and total pressure were 800 K and 1.0 MPa, respectively. Pitot pressure measurements and gas sampling for an equivalence ratio (Φ) of 0.2 showed that there were both a separated region and a supersonic core region within combustor 1 with single-stage fuel injection. For Φ = 0.4, on the other hand, a uniform subsonic region without separation was formed. One-dimensional calculation showed that two-stage fuel injection, in which the second-stage fuel was injected into combustor 2, led to formation of a large subsonic region. The maximum "thrust increment" (dF, [thrust with fuel]-[thrust without fuel]) was 37% higher than that with single-stage fuel injection, this increment being limited by occurrence of combustor-inletinteraction in both cases. Another type of two-stage fuel injection, in which both stages were within combustor 2, was conducted, and the maximum dF was found to be almost the same as that with two-stage fuel injection within both combustor 1 and combustor 2. It was also found that dF could be described as a linear function of an "effective equivalence ratio" ([total equivalence ratio] x [combustion efficiency at the combustor exit]) regardless of the fuel injection configuration. The effect on dF of the difference in total pressure loss due to the difference in fuel injection configuration was negligible because the scramjet combustor was operated in the ramjet mode. Ideal thrust was estimated by one-dimensional calculation to evaluate the "achievement factor for thrust" ([thrust obtained by experiment]/[ideal thrust obtained by one-dimensional calculation]), which was found to be 60% at most for the tested combustor under the tested conditions.

Variable Turbulent Schmidt Number Formulation for Scramjet Applications

Abstract: In high speed engines, thorough turbulent mixing of fuel and air is required to obtain high performance and high efficiency Thus, the ability to predict turbulent mixing is crucial in obtaining accurate numerical simulation of an engine and its performance Current state of the art in CFD simulation is to assume both turbulent Prandtl number and Schmidt numbers to be constants However, since the mixing of fuel and air is inversely proportional to the Schmidt number, a value of 045 for the Schmidt number will produce twice as much diffusion as that with a value of 09 Because of this, current CFD tools and models have not been able to provide the needed guidance required for the efficient design of a scramjet engine The goal of this investigation is to develop the framework needed to calculate turbulent Prandtl and Schmidt numbers as part of the solution This requires four additional equations: two for the temperature variance and its dissipation rate and two for the concentration variance and its dissipation rate In the current investigation emphasis will be placed on studying mixing without reactions For such flows, variable Prandtl number does not play a major role in determining the flow This, however, will have to be addressed when combustion is present The approach to be used is similar to that used to develop the k-zeta model In this approach, relevant equations are derived from the exact Navier-Stokes equations and each individual correlation is modeled This ensures that relevant physics is incorporated into the model equations This task has been accomplished The final set of equations have no wall or damping functions Moreover, they are tensorially consistent and Galilean invariant The derivation of the model equations is rather lengthy and thus will not be incorporated into this abstract, but will be included in the final paper As a preliminary to formulating the proposed model, the original k-zeta model with constant turbulent Prandtl and Schmidt numbers is used to model the supersonic coaxial jet mixing experiments involving He, O2 and air

Thermodynamic Analysis of Dual-Mode Scramjet Engine Operation and Performance

Abstract: Recent analytical advances in understanding the performance continuum (the thermodynamic spectrum) for air-breathing engines based on fundamental second-law considerations have clarified scramjet and ramjet operation, performance, and characteristics. Second-law based analysis is extended specifically in this work to clarify and describe the performance characteristics for dual-mode scramjet operation in the mid-speed range of flight Mach 4 to 7. This is done by a fundamental investigation of the complex but predictable interplay between heat release and irreversibilities in such an engine; results demonstrate the flow and performance character of the dual mode regime and of dual mode transition behavior. Both analytical and computational (multi-dimensional CFD) studies of sample dual-mode flow-fields are performed in order to demonstrate the second-law capability and performance and operability issues. The impact of the dual-mode regime is found to be characterized by decreasing overall irreversibility with increasing heat release, within the operability limits of the system.

Experimental Study on Combustion Mode Transition in a Scramjet with Parallel Injection

Abstract: Mode transition from weak combustion to strong combustion or vice versa in a scramjet engine is a critical phenomenon in designing such engines, because the thrust of each mode varies considerably. The mode transition is supposed to interact strongly with a so-called pseudo-shock wave or shock train. In order to control vehicles with scramjet engines, it is, therefore, essential to understand mode transition. Several studies concerning this phenomenon have been conducted and most of them used the wall fuel injection method, where the disturbance of the boundary layer due to the wall injection could not be avoided. In order to prevent this disturbance, a parallel injection method was used in this study. The experiments were conducted using the supersonic combustion facilities of

Single-stage hypersonic vehicle featuring advanced swirl combustion

Abstract: A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet/rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.

Design and Characterization of the Stanford 6 Inch Expansion Tube

Abstract: A new expansion tube facility has been constructed at Stanford University to enable research on scramjet combustor design and performance. The facility is designed to allow the duplication of combustor entrance conditions over a range of conditions covering vehicle flight Mach numbers of 4 to 9. The range can be extended to Mach 15 if the simulated pressures are allowed to vary from those of duplication. Details of the design and performance of the new facility are presented including an analysis of test repeatability, boundary layer thickness, and test gas uniformity.

Simulating Inlet Distortion Effects in a Direct-Connect Scramjet Combustor

Abstract: A special piece of hardware (called a distortion generator) was designed using computational tools to mimic the effects of inlet distortion in a direct-connect test environment. Direct-connect simulations of scramjet combustors typically use facility nozzles designed to produce uniform flow entering the test article. However, in free-jet and flight experiments, where air is ducted to the supersonic combustor through an inlet, flow entering the test article will be inherently distorted. These distortion effects can include non- uniform boundary layer thicknesses on the walls and relatively strong oblique shock waves. In this work, the design methodology for the distortion generator is described along with details of its fabrication and installation into the experimental research facility. Finally, the results of computational and experimental calibrations are presented. Results confirm that distortion characteristics anticipated in freejet and flight experiments can be effectively simulated in the direct-connect test environment. This new hardware will enable future experimental investigations aimed at understanding the effects of inlet-induced distortion on combustor operability and performance.

Modeling of plasma virtual shape control of ram/scramjet inlet and isolator

Abstract: The paper explores a concept of ram/scramjet propulsion control by energy addition and extraction in the propulsion flowpath and the reverse energy bypass concept Instead of variable geometry, the concept relies on virtual shapes created by plasma/MHD devices or by other methods (including plasma-controlled external combustion) An inherent advantage of the proposed plasma/MHD control system is its flexibility, fast response, and the absence of moving parts The fixed geometry is optimized for Mach 7 flight At Mach numbers higher than the design value, an MHD generator placed at the first compression ramp and using ionization by electron beams can restore the shock-on-lip condition, while operating in self-powered regime The magnetic field of 15-17 Tesla would be sufficient for Mach 8 flight At Mach numbers below the design value, inlet performance can be controlled by energy addition, with the power supplied by an MHD generator placed downstream of the combustor This concept is called the reverse energy bypass In one scenario, the inlet flow spillage can be reduced by Virtual Cowl – a heated region placed upstream of the cowl and slightly below it With optimally located Virtual Cowl, calculations with conservative assumption regarding power transmission losses show that the reverse bypass can increase thrust by about 10% at Mach 6 In another scenario, distributed heating of the flow upstream of the inlet throat in the ramjet regime (Mach 4-6), with the heating rate of about 63-85% of the total enthalpy flux, can bring the throat Mach number close to 1, thus making the isolator duct virtually unnecessary Although the reverse bypass system with inlet heating would reduce thrust by about 16% at Mach 5, the performance penalty at the vehicle acceleration stage can be offset by the increased efficiency during the cruise due to the absence of weight and cooling burden normally caused by the long isolator duct

Free-jet Testing of a REST Scramjet at Off-Design Conditions

Abstract: Scramjet flowpaths employing elliptical combustors have the potential to improve structural efficiency and performance relative to those using planar geometries. NASA Langley has developed a scramjet flowpath integrated into a lifting body vehicle, while transitioning from a rectangular capture area to both an elliptical throat and combustor. This Rectangular-to-Elliptical Shape Transition (REST) scramjet, has a design point of Mach 7.1, and is intended to operate with fixed-geometry between Mach 4.5 and 8.0. This paper describes initial free-jet testing of the heat-sink REST scramjet engine model at conditions simulating Mach 5.3 flight. Combustion of gaseous hydrogen fuel at equivalence ratios between 0.5 and 1.5 generated robust performance after ignition with a silane-hydrogen pilot. Facility model interactions were experienced for fuel equivalence ratios above 1.1, yet despite this, the flowpath was not unstarted by fuel addition at the Mach 5.3 test condition. Combustion tests at reduced stagnation enthalpy indicated that the engine self-started following termination of the fuel injection. Engine data is presented for the largest fuel equivalence ratio tested without facility interaction. These results indicate that this class of three-dimensional scramjet engine operates successfully at off-design conditions.

Performance of a Scramjet Combustor with Combined Normal and Tangential Fuel Injection

Abstract: The performance of a scramjet combustor with combined normal and tangential injection was experimentally investigated. Experiments were performed on a 500-mm cylindrical scramjet combustor at a freestream Mach number of 4.5, a nozzle supply pressure of 35.8 MPa, and a nozzle supply enthalpy of 5.8 MJ/kg. Hydrogen fuel was injected normally through portholes to promote combustion and tangentially through a slot to reduce viscous drag. A series of fuel injectors were used to vary the proportion of tangential to normal fuel between 45 and 100%. Reductions in the viscous drag of up to 25% were observed with the greatest reductions occurring at the lowest total equivalence ratio tested for each injector. However, the average pressure produced by combustion with combined normal and tangential injection was approximately 50% less than that produced by normal injection alone. An analysis of the change in specific impulse of the scramjet combustor indicated that the best overall performance was produced by 100% normal injection.

Strut-Based Gaseous Injection into a Supersonic Stream

Abstract: Introduction M IXING of a secondary jet with the primary supersonic stream is of great importance in many practical applications especially in scramjet combustors. Because the residence time of the high-speed flow in such combustors is only a few milliseconds, it is essential to implement mixing augmentation techniques to provide rapid and uniform mixing of fuel and air. Several injection schemes have been proposed,1−5 and it has been concluded that the vorticity is the main driving mechanism for rapid near-field mixing. The present study is conducted on the strut-based fuel injectors. The main advantage of strut-based injectors is the injection of the fuel into the core of the main flow and uniform spreading of the fuel in the lateral direction. The large-scale structures at the wake region can assist macromixing. Moreover, the shock that emanates from the leading edge of the strut is useful to enhance the mixing via the baroclinic torque mechanism.6 A further advantage of strut-based injection is the formation of a recirculation zone, which can be used for flame holding in combustion.7−9 Four types of fuel injectors are considered in the current experimental investigation, one of them being the plain strut-type injector. Diagnostic methods employed are Mie scattering combined with image processing and the time-averaged schlieren. Numerical simulation of the flowfield using the commercially available software FLUENT has also been carried out.

Scramjet Lift, Thrust and Pitching-Moment Characteristics Measured in a Shock Tunnel

Abstract: Lift, pitching moment, and thrust/drag on a supersonic combustion ramjet were measured in the T4 free-piston shock tunnel using a three-component stress-wave force balance. The scramjet model was 0.567 m long and weighed approximately 6 kg. Combustion occurred at a nozzle-supply enthalpy of 3.3 MJ/kg and nozzle-supply pressure of 32 MPa at Mach 6.6 for equivalence ratios up to 1.4. The force coefficients varied approximately linearly with equivalence ratio. The location of the center of pressure changed by 10% of the chord of the model over the range of equivalence ratios tested. Lift and pitching-moment coefficients remained constant when the nozzle-supply enthalpy was increased to 4.9 MJ/kg at an equivalence ratio of 0.8, but the thrust coefficient decreased rapidly. When the nozzle-supply pressure was reduced at a nozzle-supply enthalpy of 3.3 MJ/kg and an equivalence ratio of 0.8, the combustion-generated increment of lift and thrust was maintained at 26 MPa, but disappeared at 16 MPa. Measured lift and thrust forces agreed well with calculations made using a simplified force prediction model, but the measured pitching moment substantially exceeded predictions. Choking occurred at nozzle-supply enthalpies of less than 3.0 MJ/kg with an equivalence ratio of 0.8. The tests failed to yield a positive thrust because of the skin-friction drag that accounted for up to 50% of the fuel-off drag.

Mach 8 Ground Tests of the Hypermixer Scramjet for HyShot-IV Flight Experiment

Abstract: A hydrogen fueled scramjet engine with a wall-mounted hypermixer injector is examined at a Mach 8 flight condition in the high enthalpy shock tunnel. The hypermixer, which is previously proposed by JAXA for the use over a wide range of flight Mach numbers, uses parallel injection and generates streamwise vortices. For comparison, two injectors of normal and parallel injection without streamwise vortex generation are also examined. The engine and the injectors are full-scale models of those used in the HyShot-IV flight experiment by JAXA and the University of Queensland. The principal aims of the present ground tests are 1) to confirm the ability of the streamwise vortices to enhance mixing and combustion in a fully supersonic flow and to prevent large-scale boundary layer separation (inlet-combustor interaction), 2) to investigate the combustor working characteristics owing to the ability of the streamwise vortices, and 3) to validate the designs of the hypermixer and the combustor flowpath for the flight experiment. The results show the superior performance of the hypermixer to the other injectors in a real supersonic combustion mode at the design equivalence ratio Φ = 0.3 of HyShot-IV. In the case of hypermixer at Φ = 1.0 and 1.5, explosive combustion of the premixed fuel at the combustor exit generates a strong pressure wave, which propagates upstream followed by a combustion region and decays at the injector. Consequently a new quasi-steady supersonic combustion flowfield is formed throughout the combustor downstream of the injector. The observed pressure wave is identified as a kind of detonation wave and is suggested to propagate upstream through the streamwise vortices. We will present our main results of the ground tests on the basis of the wall pressure measurement, high-speed schlieren video, and one-dimensional flow analysis.

Mach 4 Test Results of a Dual-Flowpath, Turbine Based Combined Cycle Inlet

Abstract: An experimental study was conducted to evaluate the performance of a turbine based combined cycle (TBCC) inlet concept, consisting of a low speed turbojet inlet and high speed dual-mode scramjet inlet. The main objectives of the study were (1) to identify any interactions between the low and the high speed inlets during the mode transition phase in which both inlets are operating simultaneously and (2) to determine the effect of the low speed inlet operation on the performance of the high speed inlet. Tests were conducted at a nominal freestream Mach number of 4 using an 8 percent scale model representing a single module of a TBCC inlet. A flat plate was installed upstream of the model to produce a turbulent boundary layer which simulated the full-scale vehicle forebody boundary layer. A flowmeter/back pressure device, with remote actuation, was attached aft of the high speed inlet isolator to simulate the back pressure resulting from dual-mode scramjet combustion. Results indicate that the inlets did not interact with each other sufficiently to affect inlet operability. Flow spillage resulting from a high speed inlet unstart did not propagate far enough upstream to affect the low speed inlet. Also, a low speed inlet unstart did not cause the high speed inlet to unstart. The low speed inlet improved the performance of the high speed inlet at certain conditions by diverting a portion of the boundary layer generated on the forebody plate.

Magnetohydrodynamic energy-bypass procedure in a three-dimensional scramjet

Abstract: A generic configuration is employed to explore the kinematic and dynamic structure of the flowfield in a Mach 8 scramjet with and without magnetohydrodynamic (MHD-) bypass. Particular emphasis is placed on describing the three-dimensional interaction between the fluid and the electromagnetic field in a given plasma environment. Faraday-type operation is considered with coarsely segmented electrodes in a generator and accelerator respectively mounted on either side of a constant-area isolator/combustor element. The numerical procedure adopts a robust high-resolution technique to solve the governing equations, which include the full three-dimensional Navier-Stokes equations supplemented with electromagnetic source terms and a Poisson equation for consistency of the electric field with current continuity. Spatially variable combustion and plasma parameters are either specified or phenomenologically derived. Various three-dimensional features such as swept shock-wave boundary layer interactions have a profound impact on operation. Separated regions and vortical structures interface with the current, electric, and ponderomotive force fields to yield complex three-dimensional features that indicate that two-dimensional and inviscid analyses are inappropriate. The MHD generator shows the potential to efficiently slow down flow in the inlet, thus decreasing scramjet inlet length, and to reduce the total temperature of the flow. However, separation causes near-wall secondary eddy currents and local body force reversal, thus limiting the useful length of the generator. Accelerator operation is characterized by more prominent irreversibilities, especially near walls, where electric and body force field gradients are large. Trends in integrated pressure and viscous and magnetic force are summarized.

Scramjet Driven MHD Power Demonstration-HVEPS Program

Abstract: Under the US Air Force sponsored Hypersonic Vehicle Electric Power System (HVEPS) program the authors' organizations are collaborating on research and development of scramjet driven magnetohydrodynamic (MHD) power for an advanced high power, airborne power system. This program has been underway for the past five years with various technical tasks being addressed that have encompassed engineering investigations of a myriad of technical issues related to airborne hypersonic power system integration and operation. Under the current efforts of the program an initiative is underway to conduct an integrated scramjet-driven MHD power demonstration ground test. The demonstration test is to be conducted in the existing United Technologies Research Corporation (UTRC) scramjet test cell wherein modifications to the test cell are being made to allow installation of an in-line, direct fired MHD generator test article downstream of the test facility's scramjet combustor. This research is currently in the test preparation phase with the design and fabrication of test components including a DCW MHD generator test article and its support systems being completed and ready for installation. The MHD generator channel will be installed within the bore of a 3.0 Tesla split-coil, superconducting magnet that is being made available to support this demonstration test by NASA. The demonstration test is scheduled for conduct in the summer of 2006. This paper provides details on the HVEPS program results that have been arrived at to-date, in relation to the technical assessment of direct scramjet driven MHD power generation for specific application to a hypersonic vehicle platform. A description of the scramjet/MHD ground demonstration test is provided, in terms of the test bed design, the major test components, and its operation and general performance estimates.

Numerical Simulation of Injection of Supercritical Ethylene into Nitrogen

Abstract: A procedure for simulating the injection of supercritical ethylene into nitrogen is used to investigate aspects of the injection of supercritical fuels, considered to be an enabling technology in the design of hydrocarbons-fueled scramjet engines. The method solves the compressible Navier-Stokes equations for an ethylene/nitrogen mixture, with the thermodynamic behavior of ethylene described using the Peng-Robinson equation of state. Homogeneous equilibrium and finite-rate phase-transition models are used to describe the growth of a condensed ethylene phase in several axisymmetric and three-dimensional injector nozzles. Predictions are compared with shadowgraph and direct-lighting imaging data, mass flow rate measurements, mole-fraction and temperature measurements in the jet mixing zone, and wall pressure distributions. Qualitative trends relating to jet structure, the appearance of a condensed phase, and the effects of back pressure and injectant temperature are in good agreement with experimental results but indicate the need for improved characterization of the nozzle flow before injection and the inclusion of a better turbulence model for the jet mixing zone. For conditions where both are applicable, a nucleation/ growth phase transition model provides a similar bulk fluid response as a homogeneous equilibrium model but also yields predictions of number density and average droplet size.

Modeling Scramjet Flows with Variable Turbulent Prandtl and Schmidt Numbers

Abstract: A complete turbulence model, where the turbulent Prandtl and Schmidt numbers are calculated as part of the solution and where averages involving chemical source terms are modeled, is presented. The ability of avoiding the use of assumed or evolution Probability Distribution Functions (PDF's) results in a highly efficient algorithm for reacting flows. The predictions of the model are compared with two sets of experiments involving supersonic mixing and one involving supersonic combustion. The results demonstrate the need for consideration of turbulence/chemistry interactions in supersonic combustion. In general, good agreement with experiment is indicated.

Numerical Flow Visualization of a Single Expansion Ramp Nozzle with Hypersonic External Flow

Abstract: Numerical simulation of scramjet asymmetric nozzle flow is carried out to visualize and investigate the effects of interaction between engine exhaust and hypersonic external flow. The Single Expansion Ramp Nozzle (SERN) configuration studied here consists of flat ramp and a cowl with different combinations of ramp angle and cowl geometry. Using PARAS 3D, simulations are performed for a free stream Mach number of 6.5 that constitutes the external flow around the vehicle. Appropriate specific heats ratio has been simulated for the jet and free stream flow. External shock wave due to jet plume interaction with free stream flow, the internal barrel shock wave and the shear layer emanating from the cowl trailing edge and sidewalls are well captured. Wall static pressure distribution on the nozzle ramp for different nozzle expansion angles has been computed for both with and without side fence. Axial thrust and normal force have been evaluated by integrating the wall static pressure. Effect of cowl length variation and side fence on the SERN performance has also been studied and found to be quite significant. Based on this study, an optimum ramp angle at which the SERN generates maximum axial thrust is obtained. SERN angle of 20o was found to be optimum when the flight axis coincides with nozzle axis.

Transverse supersonic controlled swirling jet in a supersonic cross stream

Abstract: The spreading rate and mixing of a transverse jet in high-speed crossflow were modified using a swirling injector with a central control jet. The controlled supersonic swirling injector (CSSI) could be used to affect mixing both in the core and the shear layer of the jet. Rayleigh/Mie scattering from flowfield ice crystals and planar laser-induced fluorescence of the NO molecules were used to characterize penetration and mixing of the CSSI for six different cases. Instantaneous images were used to study the dynamical structures in the jet, whereas ensemble images provided information regarding the jet trajectory. Standard deviation images revealed information about the large-scale mixing/entrainment. Probability density functions were used to evaluate the probability and location of freestream, mixed, and jet fluid. They were also used to track the centerline and jet boundary on a dynamic scale. Side- (streamwise)-view images showed that the injector was capable of providing high penetration when compared to circular and swirling baseline injectors. An increase of 16% in mixing area was observed with the optimal case as compared with the other control cases. End- (spanwise)-view images show a maximum of 78 % increase in total area contained within the jet boundary for the optimal case when compared to the circular injector. Higher spanwise extent of the jet boundary was also observed with controlled cases, which could provide higher interfacial area for better mixing between the jet and the cross stream when compared to their baseline counterparts.

DES investigations of transverse injection into supersonic crossflow using a hybrid unstructured solver

Abstract: Numerical simulations of transverse injection through low-angled injector ports into a supersonic freestream are performed using a hybrid, unstructured solver. Two cases are investigated: air injected into a M=2.9 air freestream with a 25◦ injection angle, and heated helium injected into a M=4.0 air freestream with a 30◦ injection angle. Simulations were run in RANS and DES modes. A set of the DES simulations were run with an unsteady inflow boundary layer, where the perturbations are a statistically meaningful representation of a series of randomly placed hairpin eddies. This boundary condition was fed periodically into the domain, and was reused multiple times over the course of a simulation. The RANS and DES simulations are found to capture the salient features of the flow, though discrepancies with experimental data are found. While the DES simulations were found to give steady-state solutions for the flow fields, the addition of the unsteady inflow boundary layer was found to greatly impact the downstream flow field and to improve the overall agreement with experiment. In the case of the helium injection, it was found that the predicted mass fraction distributions of the DES simulations with the unsteady inflow boundary layer was far less dependent on the value of the turbulent Schmidt number. This result shows that the mixing found in DES simulation with the unsteady inflow boundary layer is a result of the large-scale turbulent motion of the flow, rather than because of the gradient diffusion term. However, the ‘box of eddies’ used to create the unsteady inflow boundary layers were not long enough to ensure that no bias was introduced into the flow field, and future simulations will be run with larger boxes. The results of the study show a great deal of promise for the use of DES simulations in conjunction with unsteady inflow boundary layers for simulation of SCRAMjet fuel injection.

Comparison of Horizontally and Vertically Launched Airbreathing and Rocket Vehicles

Abstract: A design study was performed to define and compare the parameters of horizontal- and vertical-takeoff reusable launch-vehicle systems to identify promising configurations for further developmental emphasis. The investigation considered both two-stage rockets and single-stage airbreathing ramjet/scramjet-powered vehicles, thus representing next-generation and third-generation configurations, respectively. The payload requirement for each vehicle was 20,000 Ib delivered to a 100-n mile circular Earth orbit launched easterly from Kennedy Space Center. All vehicles were first analyzed using liquid hydrogen for the entire trajectory and then reanalyzed with liquid hydrocarbon fuel for the first stage, if a rocket, or for the low-speed trajectory segment to ramjet start for the airbreathers. The vertical-takeoff airbreathing vehicles were found to have the lowest empty weights and gross takeoff weights of all of the vehicle configurations, with three-dimensional inward-turning inlets outperforming two-dimensional inlets, and the use of hydrocarbon fuel outperforming hydrogen fuel for the launch propulsive segment to ramjet start. The best horizontal takeoff vehicle is the all-hydrogen inward-turning airbreather. For the two-stage rockets, the lightest empty weight was achieved with the use of hydrocarbon fuel in the booster and hydrogen fuel in the orbiter.