Simulations of an experimental hydrogen-fueled scramjet combustor are conducted using a novel dynamic hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation (DHRL) modeling framework. The combustor has a Mach 2 core flow with a ramp fuel injector resulting in an equivalence ratio of 0.17. Three grid resolutions are obtained using local refinement by a factor of two in each direction in the fuel mixing and combustion region, and results from the three grids are used to understand the effect of grid refinement. Simulations reproduce temperature, pressure, velocity, and fuel concentrations in reasonable agreement with experimental measurements. Although heat release decreases on average, as the mesh is refined, peaks of heat release are intensified causing locally elevated temperatures. Spectral analysis of turbulence kinetic energy and heat release suggests stringent resolution requirements for reacting simulations capable of accurately resolving the effects of chemical reactions. Using the medium grid the DHRL model is compared to the improved delayed detached eddy simulation (IDDES) model and two Reynolds-averaged Navier-Stokes (RANS) models. Overall, the DHRL framework significantly outperforms other methods when compared to the experimental pressure rise. Additionally, spectral analysis suggests that the current framework is capable of accurately resolving turbulent structures at frequencies higher than IDDES. The study is the first documenting the use of DHRL for supersonic reacting flow and results suggest that it is a viable alternative to existing turbulence treatments for these types of flows.

No AccessSurvey PaperChallenges in Fuel Injection for High-Speed Propulsion SystemsJinho Lee, Kuo-Cheng Lin and Dean EklundJinho LeeNASA John H. Glenn Research Center, Cleveland, Ohio 44135, Kuo-Cheng LinTaitech, Inc., Beavercreek, Ohio 45430 and Dean EklundU.S. Air Force Research Laboratory, Wright–Patterson Air Force Base, Ohio 45433Published Online:30 Apr 2015https://doi.org/10.2514/1.J053280SectionsView Full TextPDFPDF Plus ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookTwitterLinked InRedditEmail About References [1] Weber R. J. and MacKay J. S., “Analysis of Ramjet Engines Using Supersonic Combustion,” NACA TN-4386, 1958. Google Scholar[2] Curran E. T., “Scramjet Engines: First Forty Years,” Journal of Propulsion and Power, Vol. 17, No. 6, 2001, pp. 1138–1148. doi:https://doi.org/10.2514/2.5875 JPPOEL 0748-4658 LinkGoogle Scholar[3] Bulman M. and Siebenhaar A, “The Rebirth of Round Hypersonic Propulsions,” AIAA Paper 2006-5035, July 2006. 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To make simulation-based critical decisions, decision makers need to know the level of credibility of the simulation for the intended use. Credibility depends on the accuracy of simulation, which is quantified by uncertainty. Verification and validation are processes that assess simulation accuracy. Simulation verification, propagation and quantification of uncertainties, and referent data or circumstantial evidence, are essential to establishing simulation credibility. Under the auspices of the Joint Army-Navy-NASA-Air Force (JANNAF) Interagency Propulsion Committee, a guide for assessing simulation credibility is being developed. Simulation verification, uncertainty propagation and quantification, and simulation validation are discussed. Some state-of-the-art approaches are explained and demonstrated. These approaches are anticipated to be useful to other technical communities conducting physics-based simulations.

A numerical framework is presented for CFD-based uncertainty quantification (UQ) involving propagation of combined aleatory and epistemic uncertainties to relevant output quantities, with intended application to scramjet inlet analysis. Latin hypercube sampling routines are used in combination with surrogate response surface methodology and nested Monte Carlo sampling to generate probability boxes and other UQ results based on CFD output data. For a demonstration case involving subscale Mach 6 semi-freejet testing of an inward turning inlet configuration, a total of 21 CFD simulations are performed, and probability boxes are computed for three output quantities through consideration of four different input uncertainties. Global sensitivity analysis calculations are also performed to assess relative contributions of each input uncertainty. For the current demonstration case, we find Mach number uncertainty from thermal expansion of the facility nozzle to be a dominant contributor to uncertainty in the inlet mass flow rate, while mounting angle uncertainties account for large contributions to uncertainty in wall pressure at the isolator throat.

Abstract : Experiments were performed at the Air Force Research Laboratory's Propulsion Directorate (AFRL/RZ) in Research Cell 22 (RC22). Twelve cases from the experiment were computationally analyzed and each case varied in either engine operating condition and/or combustor configuration. Initial computations were performed on all twelve cases to establish a baseline computational approach. Computations were performed on one of the cases to test sensitivity to turbulent Schmidt number, reaction rate, and grid resolution. Improvements to the baseline analysis using the results from the sensitivity analyses were extended to two additional cases. It was shown that adjustments in Schmidt number, reaction rate, or grid refinement improved the agreement with experimental data for two cases relative to the baseline results, but worsened agreement for the third case. It was left undetermined that grid refinement was a better approach to improving the baseline analysis as compared to calibrations in Schmidt number and/or reaction rate. Improvement to the grid using local refinement in regions with chemical reactions produced better results for one case and was computationally less expensive than globally refining the grid. Negligible differences were shown between results that were obtained using wall functions with Y+ value as high as 38 or results obtained using wall integration with Y+ values around one. Negligible differences were shown between periodic results that were obtained by averaging results using either a constant CFL or a constant time step. CFL-averaging a result using the constant CFL approach was 5.4 times less computationally expensive than using the constant time step approach. Computations showed that 2.53-lbm/sec. of air leaked into the exhauster housing at the exit of the combustor in RC22's test apparatus.

A technique to develop flight-like inflow conditions for direct-connect testing has been extended in the current work. This technique utilizes an expansion/contraction section upstream of the engine isolator in a direct-connect test facility to create shock waves consistent with those generated in flight. The current work improves upon the earlier work by better matching the inlet shock structures, the average conditions at the engine throat, and boundary layer properties. RANS CFD simulations have been performed for isolators downstream of a flight inlet, a direct-connect facility nozzle without distortion, and a directconnect facility nozzle with distortion. The RANS CFD simulations predict similar shock positions in all cases, however a change in the mass flux distribution at the exit plane is observed.

A PRELIMINARY numerical characterization of a U.S. Air Force Research Laboratory (AFRL) test facility was performed. Research cell 22 (RC22) is a supersonic wind-tunnel facility at AFRL. RC22 is presently simulating dual-mode combustion for circular (axisymmetric) combustors. Previous efforts in RC22 have studied combustion in rectangular flowpaths. The current effort in RC22 is to examine the benefits of axisymmetric flowpaths as compared with rectangular. For the same cross-sectional area in which circular flowpaths eliminate the challenges involved with corner flow effects, they have increased structural efficiency (heat load distribution) and have reducedweight. Circular combustors also pose challenges that involve effective fuel penetration and flame propagation. For more information regarding the rectangular experiments performed in RC22, refer to [1].

** Two engine configurations were analyzed numerically and experimentally. The first configuration had a constant area backstep and the second configuration had a tapered (divergent) combustor with no backstep. Both combustors burned gaseous ethylene fuel. The analysis used a simulated low enthalpy flight condition corresponding to Mach 3.0 flight. Numerical results showed good agreement with experimental data in terms of performance, isolator heat loss and pressure distribution. For the same fuel equivalence ratio and fuel split the tapered configuration outperformed the backstep configuration by approximately 1% in terms of stream thrust, but the backstep configuration had more isolator margin (before unstart) which allowed for higher total equivalence ratios. The backstep configuration achieved stoichiometric fueling. Model calibration via constant turbulent Schmidt number was performed for the first case in order to better match experimental results. Two different k-epsilon turbulence models were analyzed and yielded differences in pressure distribution. A conceptual design analysis involving a reduction in the number of injectors showed better fuel penetration but less circumferential coverage. Future design concepts could include changing the axial placement of the secondary injector, which could have a positive impact on operability for both configurations. Numerically, future analysis could include a variable turbulent Schmidt number which would be more representative of the flow physics and could alleviate the need for calibration in modeling reacting flows.

Using tunable diode laser absorption spectroscopy (TDLAS), water-based measurements of path-integrated static temperature, static pressure, and local velocity are made in the isolator of a direct-connect, ethylene-fueled, axisymmetric scramjet engine. Three optical lines of sight are directed across the midplane of the isolator just downstream of the facility nozzle. For the data set presented here, the scramjet engine was operated to simulate relatively low flight Mach number conditions of M ~ 3. The flight conditions were achieved using a dry, heated, but non-vitiated inflow which provided a challenging environment for the water-based direct-absorption TDLAS implementation. The TDLAS-derived temperature and velocity agree well with facility operating conditions while the derived pressure measurements highlight the need for improved pressure-broadening data. With a pressure-correction factor, the TDLAS-derived values for mass flux agree well with the measured facility mass flow. The measurements are discussed in the context of a 1-D analytical model and 3-D CFD calculations. Notably, the dynamic behavior of the velocity measurements under conditions of low isolator margin are consistent with the presence of a local shock structure predicted by the CFD computations.

This paper compares the results and the computational efficiency of a research code with that of a commercial code on the same problem. The research code (REACT-MB) is tested using its unsteady hybrid large-eddy simulation/Reynolds-averaged Navier-Stokes (LES/RANS) method as well as a more common steady-state Menter RANS method. CFD++ of Metacomp is tested on the same problem using its realizable k-ε turbulence model. Normal sonic ethylene injection through a circular injector into a Mach 2 cross-flow was simulated by each code. Time-averaged statistics of the hybrid LES/RANS computations and converged solutions from the RANS computations are compared with experimental contours of time-averaged mixture fraction. Scalability of the codes is also compared.

†‡ Numerical analysis was performed on a Dual-Mode Scramjet isolator-combustor with emphasis on the combustor. The investigation was performed to analyze overall performance and observe isolator-combustor interaction. In this investigation four studies were analyzed. It was shown by removing a flame holding cavity at the second stage of a two stage combustor, yielded negligible performance loss with a decrease in combustor complexity. A comparison study outlined marginal differences in the analysis between models performed in 2D as opposed to 3D. Another study revealed minimal performance loss by reducing the flight Mach number from Mach 3.0 to Mach 2.5 without unstarting the isolator. The final study revealed a decrease in stream thrust by introducing a less reactive fuel such as methane into an ethylene fuel mixture.

Supersonic combustion performance of a bi -component gaseous hydrocarbon fuel mixture is one of the primary aspects under investigation in the HIFiRE Flight 2 experiment. In -flight instrumentation and post -test analyses will be two key elements used to determine the combustion performance. Pre -flight computational fluid dynamics (CFD) analyses provide valuable information that can be used to optimize the placement of a constrained set of wall pressure instrumentation in the experiment. The simulations also allow pre -flight assessments of performance sensitivities leading to estimates of overall uncertainty in the determination of combustion e fficiency. Based on the pre -flight CFD results, 128 wall pressure sensors have been located throughout the isolator/combustor flowpath to minimize the error in determining the wall pressure force at Mach 8 flight conditions. Also, sensitivity analyses show that mass capture and combustor exit stream thrust are the two primary contributors to uncertainty in combustion efficiency.

Quantitative assessments of system reliability and equivalent system mass (ESM) were made for different life support architectures based primarily on International Space Station technologies. The analysis was applied to a one-year deep-space mission. System reliability was increased by adding redundancy and spares, which added to the ESM. Results were thus obtained allowing a comparison of the ESM for each architecture at equivalent levels of reliability. Although the analysis contains numerous simplifications and uncertainties, the results suggest that achieving necessary reliabilities for deep-space missions will add substantially to the life support ESM and could influence the optimal degree of life support closure. Approaches for reducing reliability impacts were investigated and are discussed.

In previous numerical simulations, a constant pressure boundary condition was imposed at the scramjet isolator exit to simulate the effect of the back-pressure valve in the experimental setup. This can, however, sometimes create a reverse flow near the outflow boundary, which in turn generates numerical instability and causes the solution to diverge. The alternative approach used here is to specify a choked nozzle in place of the backpressure valve to create a high pressure region in the scramjet isolator. The flow at the exit of the choked nozzle is supersonic, so that an extrapolation boundary condition can be applied at the outflow. The present research explores the implications of this boundary condition. Three different choked nozzle lengths were used in the prediction of the flow features within the isolator. It was found that the shorter the choked nozzle, the smaller the throat height needed to create a shock train system within the scramjet isolator. In addition, the location of the leading edge of the shock train was found to depend strongly on the height of the nozzle throat. With a longer configuration, on the other hand, higher viscous losses occur and a thicker boundary layer is formed. This creates higher Pb, back pressure at the isolator exit and thus the location of the leading edge of the shock system is less sensitive to throat height. Consequently it produces higher back pressure at the isolator exit. The leading edge of the shock train system is further upstream and closer to experimental data for the similar Pb conditions. All the numerical simulations predicted pressure oscillations along the centerline of the bottom wall, which was not observed experimentally.

An uncertainty analysis procedure was developed for the AFRL/RZAS Supersonic Combustion Research Laboratory to characterize the uncertainty associated with direct-connect-combustor performance parameters such as combustion efficiency and stream thrust. The uncertainty calculation uses a comprehensive approach (based on AGARD AR-223, ASME PTC 19.2-1998, and AIAA S-071A-1999) that includes the calibration standard accuracy, individual sensor calibrations, fundamental sensor measurements, intermediate performance variable determination, influence coefficients, and equilibrium code calculations. Knowledge of these uncertainties is critical in applying quantitative direct-connect test results in a meaningful manner in support of the combustion testing that is performed as part of our in-house development programs in hypersonic airbreathing engines. This information also guides investment strategies aimed at improving the overall confidence associated with reported engine performance values. This paper provides details of the uncertainty analysis procedure and results from a recent combustion test. Areas are identified where improved instrumentation will reduce the uncertainty associated with key performance measures. NOMENCLATURE

Tunable diode laser absorption spectroscopy (TDLAS) shows promise for in situ monitoring of oxygen mass-flux in uniform, high-speed flows as has been previously demonstrated. However, the dynamic nature of typical flows of supersonic combustors, gas turbine engines and augmenters can also lead to inhomogenities that cannot be captured by a single TDLAS measurement because of its line-of-sight nature. Instead, multiple measurements varied over several spatial locations must be made to identify the inhomogenities that develop and consequently more accurately measure the mass-flux. A key question is how to minimize the number of beam-paths required to obtain useful tomographic information about the flow. In this paper, this question is addressed using Tomographic Reconstruction by Karhunen-Loève Basis Functions (TRKB) better known as principle components analysis (PCA) as well as vector quantization. Specifically, steadystate CFD is used to form a training set of 2D profiles of oxygen and water concentrations in the isolator section of the supersonic combustion facility in research cell 18 at WrightPatterson AFB. This training set is then used to create an optimal set of basis functions that allow tomographic reconstruction to be performed with a limited number of TDLAS beam paths. This approach is also verified for the case of a laboratory scale flat flame McKenna burner. Good agreement is found between the reconstructed profiles and previously measured values for the temperature and concentration in the center flame zone as well as the ambient air.