Showing papers on "Scramjet published in 2011"

Transverse jet injection into a supersonic turbulent cross-flow

Abstract: Jet injection into a supersonic cross-flow is a challenging fluid dynamics problem in the field of aerospace engineering which has applications as part of a rocket thrust vector control system for noise control in cavities and fuel injection in scramjet combustion chambers. Several experimental and theoretical/numerical works have been conducted to explore this flow; however, there is a dearth of literature detailing the instantaneous flow which is vital to improve the efficiency of the mixing of fluids. In this paper, a sonic jet in a Mach 1.6 free-stream is studied using a finite volume Godunov type implicit large eddy simulations technique, which employs fifth-order accurate MUSCL (Monotone Upstream-centered Schemes for Conservation Laws) scheme with modified variable extrapolation and a three-stage second-order strong-stability-preserving Runge–Kutta scheme for temporal advancement. A digital filter based turbulent inflow data generation method is implemented in order to capture the physics of the sup...

CFD Analysis of the HyShot II scramjet combustor

Abstract: The development of novel air-breathing engines such as supersonic combustion ramjets (scramjets) depends on the understanding of supersonic mixing, self-ignition and combustion. These aerothermochemical processes occur together in a scramjet engine and are notoriously difficult to understand. In the present study, we aim at analyzing the HyShot II scramjet combustor mounted in the High Enthalpy Shock Tunnel Gottingen (HEG) by using Reynolds Averaged Navier Stokes (RANS) and Large Eddy Simulation (LES) models with detailed and reduced chemistry. To account for the complicated flow in the HEG facility a zonal approach is adopted in which RANS is used to simulate the flow in the HEG nozzle and test-section, providing the necessary inflow boundary conditions for more detailed RANS and LES of the reacting flow in the HyShot combustor. Comparison of predicted wall pressures and heat fluxes with experimental data show good agreement, and in particular does the LES agree well with the experimental data. The LES results are used to elucidate the flow, mixing, self-ignition and subsequent combustion processes in the combustor. The combustor flow can be separated into the mixing zone, in which turbulent mixing from the jet-in-cross flow injectors dominates, the self-ignition zone, in which self-ignition rapidly takes place, and the turbulent combustion zone, located towards the end of the combustor, in which most of the heat release and volumetric expansion takes place. Self-ignition occurs at some distance downstream of the injectors, resulting in a distinct pressure rise further downstream due to the volumetric expansion as observed in the experiments. The jet penetration is about 30% of the combustor height and the combustion efficiency is found to be around 83%.

Parametric effects on the combustion flow field of a typical strut-based scramjet combustor

Abstract: The flame-holding mechanism in hypersonic propulsion technology is the most important factor in prolonging the duration time of hypersonic vehicles. The two-dimensional coupled implicit Reynolds-averaged Navier-Stokes equations, the shear-stress transport k-ω turbulence model and the finite-rate/eddy-dissipation reaction models were used to simulate the combustion flow field of a typical strut-based scramjet combustor. We investigated the effects of the hydrogen-air reaction mechanism and fuel injection temperature and pressure on the parametric distributions in the combustor. The numerical results show qualitative agreement with the experimental data. The hydrogen-air reaction mechanism makes only a slight difference in parametric distributions along the walls of the combustor, and the expansion waves and shock waves exist in the combustor simultaneously. Furthermore, the expansion wave is formed ahead of the shock wave. A transition occurs from the shock wave to the normal shock wave when the injection pressure or temperature increases, and the reaction zone becomes broader. When the injection pressure and temperature both increase, the waves are pushed out of the combustor with subsonic flows. When the waves are generated ahead of the strut, the separation zone is formed in double near the walls of the combustor because of the interaction of the shock wave and the boundary layer. The separation zone becomes smaller and disappears with the disappearance of the shock wave. Because of the horizontal fuel injection, the vorticity is generated near the base face of the strut, and this region is the main origin for turbulent combustion.

Reduced-order modeling of turbulent reacting flows with application to ramjets and scramjets

Abstract: DOI: 10.2514/1.50272 A new engine model has been developed for applications requiring run times shorter than a few seconds, such as design optimization or control evaluation. A reduced-order model for mixing and combustion has been developed that is based on nondimensional scaling of turbulent jets in crossflow and tabulated presumed probability distribution function flamelet chemistry. The three-dimensional information from these models is then integrated across cross-sectional planes so that a one-dimensional profile of the reaction rate of each species can be established. Finally, the one-dimensional conservation equations are integrated along the downstream axial direction and the longitudinal evolution of the flow can be computed. The reduced-order model accurately simulates real-gas effects such as dissociation, recombination, and finite rate chemistry for geometries for which the main flow is nearly onedimensional. Thus, this approach may be applied to any flowpath in which this is the case; ramjets, scramjets, and rockets are good candidates. Comparisons to computational fluid dynamics solutions and experimental data were conducted to determine the validity of this approach. I. Introduction T HIS work addresses the need for an improved control-oriented model of a dual-mode ramjet/scramjet propulsion system. Improvements are needed to include more realistic estimates of the losses of the propulsion efficiency due to shock wave interactions in the inlet, as well as due to gas dissociation and incomplete combustion in the combustor section. One problem is that previous lowerorder propulsion models [1–3] do not include the losses due to multiple shock interactions, gas dissociation, and incomplete combustioncausedby finiteratechemistry.Thisisaseriousproblem, because the main advantage of ascramjet engine over a ramjet isthat the scramjet reduces losses due to internal shock waves and gas dissociation [4]. That is, the scramjet eliminates the need for strong internalshockwavestodeceleratethegastosubsonicconditionsand maintains lower static temperatures than a ramjet, which reduces the dissociation losses. The present effort addresses previous shortcomings by including both of these types of losses into a code called MASIV. MASIV consists of several reduced-order models (ROMs). One is an inlet ROM that computes losses due to multiple shock/ expansion wave interactions; this ROM is described elsewhere [5]. The other is a fuel/air mixing/combustion ROM that is the focus of the present paper. MASIV has been incorporated into a larger hypersonic vehicle (HSV) code, which is available without charge and without International Traffic in Arms Regulations restrictions. Sincecomputational fluiddynamics(CFD)codestakemanyhours to reach solutions for reacting flows, they are difficult to apply to problems in which a large number of solutions are required. A tool that can solve these configurations in a short time to acceptable accuracyishighlydesirableforcontrolanddesignapplications,such

Simultaneous measurements of multiple flow parameters for scramjet characterization using tunable diode-laser sensors

Abstract: This paper reports the simultaneous measurements of multiple flow parameters in a scramjet facility operating at a nominal Mach number of 2.5 using a sensing system based on tunable diode-laser absorption spectroscopy (TDLAS). The TDLAS system measures velocity, temperature, and water vapor partial pressure at three different locations of the scramjet: the inlet, the combustion region near the flame stabilization cavity, and the exit of the combustor. These measurements enable the determination of the variation of the Mach number and the combustion mode in the scramjet engine, which are critical for evaluating the combustion efficiency and optimizing engine performance. The results obtained in this work clearly demonstrated the applicability of TDLAS sensors in harsh and high-speed environments. The TDLAS system, due to its unique virtues, is expected to play an important role in the development of scramjet engines.

Experimental Study of Test-Medium Vitiation Effects on Dual-Mode Scramjet Performance

Abstract: An experimental study was performed to characterize the effects of vitiation due to combustion-air preheating on dual-mode scramjet combustion.Major combustion vitiation species (H2O andCO2)were added to the freestreamof an electrical-resistance-heated, direct-connect facility simulating Mach 5 flight enthalpy. With clean, dry air, the combustor operated in the supersonic mode at fuel equivalence ratios below 0.22, and in the subsonic mode for equivalence ratios above 0.26. Hysteresis was observed in the dual-mode transition region between 0.22 and 0.26, as the mode of combustion was dependent on whether the fuel rate was increasing or decreasing. Adding increasing amounts of water vapor and carbon dioxide to the freestream decreased combustor pressures by 10 to 30% for the same fuel equivalence ratio. Vitiation also caused transition between supersonic and subsonic combustion to occur at a higher fuel equivalence ratio thanwith clean air. This work represents the first direct evaluation of the effect of testmedium vitiation on dual-mode scramjet combustion atMach 5 enthalpy simulation in the same facility. The results indicate the importance of accounting for test-medium vitiation when extrapolating from ground-testing to flight, particularly in the dual-mode transition region between subsonic and supersonic combustion regimes.

Investigation on the flameholding mechanisms in supersonic flows: backward-facing step and cavity

Abstract: As effective devices to extend the fuel residence time in supersonic flow and prolong the duration time for hypersonic vehicles cruising in the near-space with power, the backward-facing step and the cavity are widely employed in hypersonic airbreathing propulsive systems as flameholders. The two-dimensional coupled implicit RANS equations, the standard k-e turbulence model, and the finite-rate/eddy-dissipation reaction model have been used to generate the flow field structures in the scramjet combustors with the backward-facing step and the cavity flameholders. The flameholding mechanism in the combustor has been investigated by comparing the flow field in the corner region of the backward-facing step with that around the cavity flameholder. The obtained results show that the numerical simulation results are in good agreement with the experimental data, and the different grid scales make only a slight difference to the numerical results. The vortices formed in the corner region of the backward-facing step, in the cavity and upstream of the fuel injector make a large difference to the enhancement of the mixing between the fuel and the free airstream, and they can prolong the residence time of the mixture and improve the combustion efficiency in the supersonic flow. The size of the recirculation zone in the scramjet combustor partially depends on the distance between the injection and the leading edge of the cavity. Further, the shock waves in the scramjet combustor with the cavity flameholder are much stronger than those that occur in the scramjet combustor with the backward-facing step, and this causes a large increase in the static pressure along the walls of the combustor.

Investigation on the flameholding mechanisms in supersonic flows: backward-facing step and cavity flameholder

Abstract: As effective devices to extend the fuel residence time in supersonic flow and prolong the duration time for hypersonic vehicles cruising in the near-space with power, the backward-facing step and the cavity are widely employed in hypersonic airbreathing propulsive systems as flameholders. The two-dimensional coupled implicit RANS equations, the standard k-e turbulence model, and the finite-rate/eddy-dissipation reaction model have been used to generate the flow field structures in the scramjet combustors with the backward-facing step and the cavity flameholders. The flameholding mechanism in the combustor has been investigated by comparing the flow field in the corner region of the backward-facing step with that around the cavity flameholder. The obtained results show that the numerical simulation results are in good agreement with the experimental data, and the different grid scales make only a slight difference to the numerical results. The vortices formed in the corner region of the backward-facing step, in the cavity and upstream of the fuel injector make a large difference to the enhancement of the mixing between the fuel and the free airstream, and they can prolong the residence time of the mixture and improve the combustion efficiency in the supersonic flow. The size of the recirculation zone in the scramjet combustor partially depends on the distance between the injection and the leading edge of the cavity. Further, the shock waves in the scramjet combustor with the cavity flameholder are much stronger than those that occur in the scramjet combustor with the backward-facing step, and this causes a large increase in the static pressure along the walls of the combustor. Graphical Abstract text

Experimental Investigations on Transpiration Cooling for Scramjet Applications Using Different Coolants

Abstract: The extremely high heat loads within a scramjet combustor require the use of high-temperature materials combined with efficient cooling concepts. A promising technique is the application of transpiration cooling to ceramic matrix composite materials. A supersonic hot-gas-flow test facility is used to investigate this cooling method. The carbon/carbon samples tested have porosities of about e = 11%. The airflow is electrically heated up to 1120 K total temperature with a total pressure of ≈3 bar and is accelerated to a Mach number of 2.1 within the test channel. Air, argon, and helium are used as coolants for blowing ratios from 0 to 1 %. The surface temperature of the porous wall is measured via thermocouples and infrared thermography. Pressure and mass-flow measurements are used to analyze the throughflow characteristics of the porous materials at various temperature conditions. An approach based upon simplified analytical models is presented to analyze the experimental data of throughflow behavior and cooling efficiency. The simplified thermal model is used to analyze the effect of fluid property variations with temperature on pressure loss for different coolants and shows good agreement with the experimental data.

Combustor Operability and Performance Verification for HIFiRE Flight 2

Abstract: As part of the Hypersonic International Flight Research Experimentation (HIFiRE) Direct-Connect Rig (HDCR) test and analysis activity, three-dimensional computational fluid dynamics (CFD) simulations were performed using two Reynolds-Averaged Navier Stokes solvers. Measurements obtained from ground testing in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) were used to specify inflow conditions for the simulations and combustor data from four representative tests were used as benchmarks. Test cases at simulated flight enthalpies of Mach 5.84, 6.5, 7.5, and 8.0 were analyzed. Modeling parameters (e.g., turbulent Schmidt number and compressibility treatment) were tuned such that the CFD results closely matched the experimental results. The tuned modeling parameters were used to establish a standard practice in HIFiRE combustor analysis. Combustor performance and operating mode were examined and were found to meet or exceed the objectives of the HIFiRE Flight 2 experiment. In addition, the calibrated CFD tools were then applied to make predictions of combustor operation and performance for the flight configuration and to aid in understanding the impacts of ground and flight uncertainties on combustor operation.

SCRAMSPACE : Scramjet-based access-to-space systems

Abstract: Scramjet-based launch systems offer considerable promise for safe, reliable and economical access to space. Through both flight and ground tests, leveraging Australia's world leadership in scramjet R and D, the SCRAMSPACE project will answer key scientific and technological questions and build an industry-ready talent pool for a future Australian scramjet-based access-to-space industry.

Free-piston driver optimisation for simulation of high Mach number scramjet flow conditions

Abstract: The University of Queensland (UQ) is currently developing high Mach number, high total pressure scramjet flow conditions in its X2 and X3 expansion tube facilities. These conditions involve shock-processing a high-density air test gas followed by its unsteady expansion into a low-pressure acceleration tube. This relatively slow shock-processing requires the driver to supply high pressure gas for a significantly greater duration than normally required for superorbital flow conditions. One technique to extend the duration is to operate a tuned free-piston driver. For X2, this involves the use of a very light piston at high speeds so that, following diaphragm rupture, the piston displacement substitutes for vented driver gas, thus maintaining driver pressure much longer. However, this presents challenges in terms of higher piston loading and also safely stopping the piston. This article discusses the tuned driver concept, the design of a very lightweight but highly stressed piston, and details the successful development of a new set of tuned free-piston driver conditions for X2.

Combined Cycle Propulsion: Aerojet Innovations for Practical Hypersonic Vehicles

Abstract: Sustained hypersonic (M>5) atmospheric flight is best performed with a supersonic combustion ramjet (Scramjet). Scramjet-powered flight with durations over 100 seconds (X51) has now been demonstrated. A Scramjet-powered vehicle requires a “boost” to its ramjet takeover (RTO) speed where it can provide adequate acceleration to its cruise speed. All Scramjet powered flights to date have used a solid rocket booster motor that was dropped just before RTO. Although this is practical for demonstration flights, it is unaffordable for an operational vehicle, particularly when the vehicle is to be reused. Reusable hypersonic vehicles require a synergistically integrated self boosting capability. In this paper, we describe the issues such propulsion systems must address and discuss various Aerojet solutions to make them practical.

Experimental Characterization of Hydrocarbon-Fueled, Axisymmetric Scramjet Combustor Flowpaths

Abstract: A summary is presented of recent experimental efforts at the Air Force Research Laboratory (AFRL/RZAS) related to the development, calibration, and evaluation of axisymmetric scramjet combustor flowpaths. The flowfields exiting the facility nozzles and isolators in the direct-connect rig were characterized using a radially traversing probe at conditions that simulated flight Mach numbers between 3 and 5. The performance characteristics of two isolator configurations were compared using a throttling, backpressure valve. The fully divergent isolator was found to sustain a slightly higher pressure ratio and was used for all of the subsequent combustor evaluations. Two cavity-based combustor configurations with equivalent exit-to-inlet area ratios were evaluated. One configuration was continuously divergent, whereas the other incorporated a step flameholder downstream of the cavity. Heated ethylene was used to fuel the combustors over an equivalence ratio range 0.2 <  < 1.0. The wall pressure distribution in the divergent combustor exhibited a single pressure peak whereas the step-combustor pressure profiles exhibited a secondary pressure rise downstream of the sudden expansion. The shape and magnitude of this secondary pressure rise was sensitive to the overall equivalence ratio and the fuel distribution between the primary and secondary injectors. The step combustor exhibited a wider operating margin and higher combustion efficiency because of the area relief and the second flameholding zone. The effects of varying fuel distribution between the primary and secondary fuel-injection sites were also explored. Stream thrust based on Pitotpressure measurements at the combustor exit compares very favorably with load-cell measurements, and an in-stream imaging probe provided qualitative information related to combustor operation and the effects of fuel distribution in the combustor.

CFD Analysis of Mixing and Combustion of a Scramjet Combustor with a Planer Strut Injector

Abstract: As one of the most promising propulsive systems in the future, the scramjet engine has drawn the attention of many researchers. The two-dimensional coupled implicit NS equations, the standard k-e turbulence model and the finite-rate/eddy-dissipation reaction model have been applied to numerically simulate the flow field of the hydrogen fueled scramjet combustor with a planer strut flame holder under two different working conditions, namely, cold flow and engine ignition. The obtained results show that the numerical method used in this paper is suitable to simulate the flow field of the scramjet combustor. The static pressure distribution along the top and bottom walls for the case under the condition of engine ignition is much higher than that for the case under the condition of cold flow. There are three clear pressure rises on the top and bottom walls of the scramjet combustor. The eddy generated in the strut acts as a flame holder in the combustor, and it can prolong the residence time of the mixture in the supersonic flow.

HIFiRE Direct-Connect Rig (HDCR) Phase I Scramjet Test Results from the NASA Langley Arc-Heated Scramjet Test Facility

Abstract: A series of hydrocarbon-fueled direct-connect scramjet ground tests has been completed in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) at simulated Mach 8 flight conditions. These experiments were part of an initial test phase to support Flight 2 of the Hypersonic International Flight Research Experimentation (HIFiRE) Program. In this flight experiment, a hydrocarbon-fueled scramjet is intended to demonstrate transition from dual-mode to scramjet-mode operation and verify the scramjet performance prediction and design tools A performance goal is the achievement of a combusted fuel equivalence ratio greater than 0.7 while in scramjet mode. The ground test rig, designated the HIFiRE Direct Connect Rig (HDCR), is a full-scale, heat sink test article that duplicates both the flowpath lines and a majority of the instrumentation layout of the isolator and combustor portion of the flight test hardware. The primary objectives of the HDCR Phase I tests were to verify the operability of the HIFiRE isolator/combustor across the simulated Mach 6-8 flight regime and to establish a fuel distribution schedule to ensure a successful mode transition. Both of these objectives were achieved prior to the HiFIRE Flight 2 payload Critical Design Review. Mach 8 ground test results are presented in this report, including flowpath surface pressure distributions that demonstrate the operation of the flowpath in scramjet-mode over a small range of test conditions around the nominal Mach 8 simulation, as well as over a range of fuel equivalence ratios. Flowpath analysis using ground test data is presented elsewhere; however, limited comparisons with analytical predictions suggest that both scramjet-mode operation and the combustion performance objective are achieved at Mach 8 conditions.

An Experimental Investigation of Steady and Unsteady Combustion Phenomena in the HyShot II Combustor

Abstract: A series of experiments has been carried out in the HEG facility (High Enthalpy Shock Tunnel, Gottingen) to obtain detailed measurements on the HyShotII scramjet configuration under both steady and unsteady combustion conditions. Standard pressure measurements are performed, but a main focus of this campaign is the use of optical and visualization techniques: high-speed visualizations of OH* chemiluminescence together with pulsed-diode laser Schlieren imaging are employed to gain information about the approximate flame location as well as the interaction between the flow features and combustion characteristics. In particular, this allows an unprecedented level of insight into the transient combustioninduced phenomena present in the combustion chamber at high equivalence ratios. Measurements are compared with computational data obtained from the DLR TAU code. I. Introduction

Performance Analysis of Variable-Geometry Scramjet Inlets Using a Low-Order Model

Abstract: Scramjet vehicles, especially those used as part of an orbital launch system, must operate over a wide range of flight conditions. One component that has difficulty accommodating a range of Mach numbers is the inlet. In this article, only two-dimensional-type scramjet inlets are considered. Such an inlet with fixed geometry can be designed for a single Mach number (using approximately the shock-on-lip configuration) or a range of Mach numbers. However, the performance of the inlet tends to degrade as the size of the Mach number range increases. One method to improve this performance is to use a variable-geometry cowl. Three cowl motions are considered in this paper: moving the whole cowl up and down, moving the whole cowl forward and backward, and rotating the cowl lip. A low-order model designed for control-oriented applications is used to simulate wave interactions. The model is used to evaluate the benefits of each type of variable geometry, and an inlet designed for a wide range of Mach numbers is presented.

Effect of cavity location on combustion flow field of integrated hypersonic vehicle in near space

Abstract: The cavity has been widely employed as the flame holder to prolong the residence time of fuel in supersonic flows since it improves the combustion efficiency in the scramjet combustor, and also imposes additional drag on the engine. In this paper, the two-dimensional coupled implicit Reynolds Average Navier–Stokes equations, the RNG k–e turbulence model and the finite-rate/eddy-dissipation reaction model have been employed to numerically simulate the combustion flow field of an integrated hypersonic vehicle. The effect of cavity location on the combustion flow field of the vehicle has been investigated, and the fuel, namely hydrogen, was injected upstream of the cavity on the walls of the first stage combustor. The obtained results show that the viscous lift force, drag force and pitching moment of the vehicle are nearly unchanged by varying the cavity location over the location range and designs considered in this article, namely the configurations with single cavity, double cavities in tandem and double cavities in parallel. The variation of the fuel injection strategy affects the separation of the boundary layer, and the viscous effect on the drag force of the vehicle is remarkable, but the viscous effects on the lift force and the pitching moment are both small and they can be neglected in the design process of hypersonic vehicles. In addition to varying the location of the cavities, three fuel injection configurations were considered. It was found that one particular case can restrict the inlet unstart for the scramjet engine.

Parametric Ideal Scramjet Cycle Analysis

Abstract: Parametric cycle analyses for the ideal ramjet, ideal turbojet, and ideal turbofan engines are well known and documented. The parametric mathematical descriptions of these ideal propulsion systems are useful for understanding the advantages and useful operation conditions when comparing these engines with one another and with mission requirements. It is also known that the scramjet engine is superior in producing specific thrust over these other engineswhen operating at hypersonicMachnumbers. The purpose of this work is to present a parametric cycle analysis for the ideal scramjet. This permits the description of the ideal scramjet via simple algebraic equations similar to that for the ideal ramjet, ideal turbojet, and ideal turbofan engines.

HIFIRE Flight 2 Overview and Status Update 2011

Abstract: A collaborative international effort, the Hypersonic International Flight Research Experimentation (HIFiRE) Program aims to study basic hypersonic phenomena through flight experimentation. HIFiRE Flight 2 teams the United States Air Force Research Lab (AFRL), NASA, and the Australian Defence Science and Technology Organisation (DSTO). Flight 2 will develop an alternative test technique for acquiring high enthalpy scramjet flight test data, allowing exploration of accelerating hydrocarbon-fueled scramjet performance and dual-to-scram mode transition up to and beyond Mach 8 flight. The generic scramjet flowpath is research quality and the test fuel is a simple surrogate for an endothermically cracked liquid hydrocarbon fuel. HIFiRE Flight 2 will be a first of its kind in contribution to scramjets. The HIFiRE program builds upon the HyShot and HYCAUSE programs and aims to leverage the low-cost flight test technique developed in those programs. It will explore suppressed trajectories of a sounding rocket propelled test article and their utility in studying ramjet-scramjet mode transition and flame extinction limits research. This paper describes the overall scramjet flight test experiment mission goals and objectives, flight test approach and strategy, ground test and analysis summary, development status and project schedule. A successful launch and operation will present to the scramjet community valuable flight test data in addition to a new tool, and vehicle, with which to explore high enthalpy scramjet technologies.

Cowl and Cavity Effects on Mixing and Combustion in Scramjet Engines

Abstract: To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the T4 free-piston shock tunnel. The test model had a rectangular intake, which compressed the freestream flow through a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of 31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40. Furthermore, the cavity flame holder and the W-shaped cowl showed improved pressure distribution due to greater reaction intensity. With the aid of numerical analysis, the cavity and the W-shaped cowl are shown to be effective in fuel–air mixing.

Uncertainty quantification and error estimation in scramjet simulation

Abstract: The numerical prediction of scramjet in-flight performance is a landmark example in which current simulation capability is overwhelmed by abundant uncertainty and error. The aim of this work is to develop a decision-making tool for balancing the available computational resources in order to equally reduce the effects of all sources of uncertainty and error below a confidence threshold. To that end, a nested uncertainty quantification and error estimation loop is proposed that balances aleatoric uncertainty, epistemic uncertainty, and numerical error in an efficient way. Application to a nozzle flow problem shows a reduction of the confidence interval by three orders of magnitude. The framework applied to the HyShot II scramjet flight experiment validation simulation indicates that the epistemic uncertainty in the RANS turbulence model is the dominating contribution to the confidence interval.

Experimental investigation of combustion mechanisms of kerosene-fueled scramjet engines with double-cavity flameholders

Abstract: A scramjet combustor with double cavitybased flameholders was experimentally studied in a directconnected test bed with the inflow conditions of M = 2.64, P t = 1.84MPa, T t = 1 300 K. Successful ignition and self-sustained combustion with room temperature kerosene was achieved using pilot hydrogen, and kerosene was vertically injected into the combustor through 4×ϕ0.5mm holes mounted on the wall. For different equivalence ratios and different injection schemes with both tandem cavities and parallel cavities, flow fields were obtained and compared using a high speed camera and a Schlieren system. Results revealed that the combustor inside the flow field was greatly influenced by the cavity installation scheme, cavities in tandem easily to form a single side flame distribution, and cavities in parallel are more likely to form a joint flame, forming a choked combustion mode. The supersonic combustion flame was a kind of diffusion flame and there were two kinds of combustion modes. In the unchoked combustion mode, both subsonic and supersonic combustion regions existed. While in the choked mode, the combustion region was fully subsonic with strong shock propagating upstream. Results also showed that there was a balance point between the boundary separation and shock enhanced combustion, depending on the intensity of heat release.

Characterizing the operability limits of the HyShot II scramjet through RANS simulations

Abstract: Experimental and flight data for hypersonic air-breathing vehicles are both difficult and extremely expensive to obtain, motivating the use of computational tools to enhance our understanding of the complex physics involved. One of the major difficulties in simulating this regime is the interaction between combustion and turbulence, both of which are intrinsically complex processes. This work represents a first attempt at addressing assumptions introduced by physical models representing the turbulent reacting flow on the resulting predictions of the scramjet performance. A combustion model for high-speed flows is introduced and tested for the HyShot II vehicle. A reduced order chemistry model is then derived to investigate the effect of certain chemistry modeling assumptions within the combustion model. These models are used to investigate the unstart of the engine due to thermal choking by increasing the fuel flow rate. It is shown that an abrupt change occurs where a normal shock forms and moves upstream accompanied by a large region of subsonic flow. Additionally scalar metrics are described which are used as early indicators of unstart, to formulate safe operating limits for the scramjet engine.

Demonstration of Capability of Water Flux Measurement in a Scramjet Combustor using Tunable Diode Laser Absorption Tomography and Stereoscopic PIV

Abstract: Tunable Diode Laser Absorption Tomography (TDLAT) is a non-intrusive measurement technique for determining two-dimensional spatially resolved distributions of temperature and species concentration in high enthalpy flows. TDLAT combines infrared laser absorption spectroscopy with tomographic image reconstruction. The TDLAT technique has been implemented at the University of Virginia's Aerospace Research Laboratory supersonic combustion tunnel. Spatially resolved temperature and water vapor concentration measurements at the exit of the supersonic combustion tunnel have been obtained using the TDLAT technique. The tomographic reconstructions at the exit plane are presented. Water vapor concentration measurements from TDLAT are combined with velocity measurements obtained by Stereoscopic PIV to provide direct measurement of water vapor flux at the exit of the supersonic combustion tunnel. Comparison of this value to the known water vapor injected provides a demonstration of the capability of the TDLAT/SPIV technique. Such a measurement of water flux with combustion will enable combustion efficiency to be evaluated. A measurement of a high temperature non-reacting case for a water vitiation level of 12% is presented.