A Study of Ram Combustor Performance Using 1D and 3D Numerical Simulations

S. Yungster Ohio Aerospace Institute/NASA Glenn Research Center, Brook Park OH, 44135, USA

A. Suresh

M. E. M. Stewart QSS/NASA Glenn Research Center, Brook Park OH, 44135, USA

J. Lee

NASA Glenn Research Center, Brook Park, OH 44135, USA

I. Abstract Detailed three-dimensional and one-dimensional simulations in support of the Direct Connect Combustor experiment (DCC) are

presented. The one-dimensional approach, based on empirical isolator and combustor models, solves area averaged one-dimensional flow in the combustor with equilibrium chemistry. The three-dimensional simulations are performed using a RANS solver with a finite-rate chemistry model for ethylene. Isolator unstart and ignition were two major difficulties encountered in the three-dimensional simulations. Despite these issues, reasonable agreement exists between the one-dimensional and three-dimensional computations.

II. Introduction This paper describes numerical simulations of varying levels of fidelity, performed in support of the NASA ISTAR Direct Connect Combustor Experiment (DCC) at Glenn Research Center. ISTAR (Integrated System Test of an Airbreathing Rocket) is an engine system/integrated vehicle development project focused on a Rocket Based Combined Cycle (RBCC) engine. The ISTAR mission is to accelerate a hydrocarbon fueled vehicle through subsonic, supersonic and hypersonic speed ranges. An RBCC engine combines rockets at lower speeds with airbreathing RAM/SCRAM operation at higher speeds. Although rocket engines alone routinely operate in this regime, RBCC engines promise lower cost access to space by exploiting the higher specific impulse of airbreathing hypersonic propulsion. Figure 1 shows two views of the ISTAR concept vehicle and engine system.

Figure 1: Two views of the ISTAR concept vehicle and engine system The DCC experiment is an experimental program to validate the ISTAR engine’s combustion performance at important points in the flight trajectory. As a large-scale combustor rig, it avoids combustion scaling challenges (i.e. flame holding) and compromised testing conditions. Further, the rig’s scale reinforces the primary goals of this combustor experiment, namely, to provide quantitative data to address technology gaps and to enhance understanding of endothermic hydrocarbon-fueled dual-mode combustors. While obtaining the overall combustion efficiency of the dual-mode burner is an important part of the proposed experimental study,

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it is just as critical to show that adequate flame holding behavior can be maintained during combustion system mode changes in the flight mission. It is also essential to achieve an understanding of the operability of the burner to evaluate system stability. Hence, the combustor performance will be evaluated as a function of several parameters, such as injector locations, fuel splits and properties.

Initially, the experimental investigation will focus on the low flight enthalpy Mach 3.5 condition. Some of the specific objectives are:

a) to study the impact of injector locations and fuel splits on the burner pressure characteristics, associated isolator behavior, thermal throat location, and combustion efficiency;

b) to study the ability to hold a stable flame via piloted base regions and fuel flow ranges needed; c) to study the sensitivity of the flame in the combustor to entry condition perturbations and fuel state; d) to study the thermal environment of the burner under various flight conditions; e) to study if the isolator can retain the combustion-induced backpressure, avoiding isolator unstart.

Further details of the test program and its objectives can be found in Reference 1.

Figure 2 shows the dominant physical phenomena of the combustor rig. The key physical features include: 1. sufficiently high combustion pressure will establish a subsonic region in the combustor; this low speed

flow increases fuel penetration, residency time, and combustor efficiency. 2. a series of shocks (a shock train) is established in the isolator duct which accommodates the adverse

pressure gradient between the inlet and the combustor. 3. combustion heat release must establish a thermal choke and supersonic flow in the nozzle. 4. due to the nozzle’s changing area, the position of the thermal choke in the nozzle strongly influences the

pressure in the combustor. 5. the recirculating flow in the base region plus the injected fuel provide a stable pilot region.

M>1M>1 Side View M<1

Figure 2: Side and Planform view schematics of dominant Combustor physics.

The main rationale of these numerical simulations was to help validate the experiment’s design, and then use the DCC experimental results to help validate the analytical tools needed to design a flight engine. In particular, for a baseline geometry and injector parameters (location), the goal of this analytical effort was to predict critical quantities such as the combustion backpressure, the thermal throat location, and combustion efficiency. Further, since the simulation conditions are the low enthalpy Mach 3.5 condition, ignition and flame blow out

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are important engineering issues. Despite the fidelity of three-dimensional Navier-Stokes solutions with the existing finite-rate chemistry models, prior experience2 suggests that these simulations cannot be used to estimate flame blow out limits or ignition characteristics. At best, these solutions are an optimistic picture of the combustion process.

II. Three-Dimensional Simulations

The steady, chemically reacting flow within the inlet, isolator duct, and combustor is simulated with VULCAN3 which solves the three-dimensional Reynolds Averaged Navier-Stokes (RANS) equations with a finite-rate chemistry model. Turbulence effects are simulated with a κ-ω turbulence closure model; wall functions are used so the grid resolution of the boundary layer is y+

max< 500. Here, only laminar finite-rate chemistry effects were considered. The grid is a composite (non-overlapping) structured grid containing 1.63×106 points split into 19 blocks. The geometry is shown schematically in Figure 3.

Cascade Injectors

Body and Cowl Injectors

Isolator Strut

Base pilot injectors

Figure 3: Geometry and grid layout for the DCC simulation. The injectors are modeled as rectangular slots with the correct area and mass flow. A sonic or supersonic boundary condition was used on all injectors. Wall functions were used in the simulations so that there was no need to resolve the near wall region. The inflow and outflow boundaries are both supersonic with standard boundary conditions applied. Presently, all solid walls are imposed with adiabatic, no-slip conditions, although these could be changed to isothermal walls to get estimates of the heat flux. A coupled fluid-solid solution is also feasible along the lines described in Reference 4. The modeling of combustion is the pivotal element of the simulations presented. Although the DCC experiment will burn JP-7 fuel, in this simulation the JP-7 fuel is assumed to be thermally and catalytically cracked into gaseous ethylene. Combustion of ethylene is modeled with a 6-species, 3-equation finite-rate kinetics model3, 5. Another option, not presented here, is a mixed-is-burned calculation with JP-7 fuel using the Eddy Dissipation Concept (EDC) of Magnussen and Hjertager6 for which a kinetic reaction mechanism is not required. These calculations are very similar to computations initially carried out for the ISTAR flight vehicle and reported in Reference 7. The notable difference is the absence of a forebody, inlet and rockets.

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These three-dimensional computations require considerable computational resources compared to one-dimensional calculations. Typical run times on a Linux cluster (19 processors) are approximately 500 iterations per hour. However, these three-dimensional calculations are more detailed, more predictive, and less dependent on empirical data and correlations because they include more of the detailed physics and three-dimensional geometry. Indeed, turbulence and chemistry modeling are the largest sources of uncertainty in the simulation results.

III. One-Dimensional Simulations A quasi one-dimensional computational code was developed to analyze ramjet/scramjet flows in a quick and inexpensive manner. It is composed of two parts: an isolator model, and a combustor model. The models are applicable to any fuel/oxidizer combination, assume chemical equilibrium, and can be used for both subsonic and supersonic conditions. The code is capable of modeling dual mode operation without the need for any special treatment at the thermal throat. Models can be included for skin friction and heat transfer.

The isolator is a section of the flowpath whose function is to de-couple the inlet from the pressure rise that occurs in the combustor. The isolator sustains the pressure rise through a shock train interaction with the boundary layer. There are several isolator models that have been proposed. In the present study, the Sullins-McLafferty isolator model8 is used. This model is applicable to ducts having rectangular or circular cross-sections. A quasi one-dimensional CFD approximation is used to model the flow through the combustor. Combustion is modeled by a prescribed distribution of heat release along the combustor. Only the shape of the heat release profile as a function of distance is required as input to the code. Equilibrium chemistry is used to model the combustion process. The quasi one-dimensional Euler equations for a gas in chemical equilibrium are solved using a second-order total variation diminishing (TVD) McCormack scheme, and using the LSENS chemical kinetics code9 to compute the equilibrium composition. Complete details about the code are given in Reference 10. This code has been validated by calculating the ramjet flowfield in the Injector Characterization Rig (ICR) at NASA Langley Research Center and comparing the results with experiments and with the HyCAD numerical code. Figure 4 shows pressure and Mach number profiles. Good agreement between the experimental and numerical data was obtained. These results indicate that the present quasi one-dimensional code can model reasonably well the flow in a ramjet/scramjet engine with minimal computational cost. This approximate method complements the more detailed three-dimensional CFD studies. Preliminary comparisons between the one-dimensional and three-dimensional calculations for the DCC experiment are presented below.

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Figure 4: Comparison of the one-dimensional model with experiment and other models for the ICR test

data. Another possible one-dimensional approach that we mention and are pursuing is to reduce the steady one-dimensional Euler equations with varying area, heat addition, mass addition and friction to a single ordinary differential equation (ODE) for the Mach number. For ideal gases with a constant gamma this approach can be found in Reference 11. Required inputs are the area distribution and the total temperature distribution in the duct. This equation is singular at sonic points and in particular at the thermal choke. However, the location of the thermal choke can be obtained in closed form without solving for the whole flow field. Once the thermal choke location is determined, the ODE for Mach number is integrated by marching upstream and downstream of the thermal choke using standard methods. All other quantities are obtained from the equations and the boundary conditions.

Since the location of the thermal choke is a crucial design criterion for RAM combustors, this approach is well suited for optimization studies. For example, a simple optimization problem that is of interest is as follows: Given a fixed 1D total temperature distribution in the combustor, determine the optimum area distribution for maximum thrust. This simplified approach assumes that small changes in the area distribution do not affect the heat release, an assumption that is clearly incorrect.

IV. Preliminary Results Figure 5 shows results from the three-dimensional simulation. These solutions were obtained with the finite rate chemistry model, and with the activation energies lowered to facilitate ignition (see discussion below). Although not visible, there is significant separation on the top wall of the isolator in both ducts. Figure 5(b) shows iso-surfaces of Mach number equal to 1 (sonic surface). The colored areas are supersonic regions while the empty areas are subsonic. The presence of a shock train in the isolator duct can be inferred from Figure 5(b). In addition, a well defined thermal choke can be seen just downstream of the body and cowl injectors. Figure 6 shows the combustion region in greater detail. In particular the cascade and body/cowl fuel injectors (denoted by plumes), reverse flow regions, and combustor geometry are visible.

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Figure 5: (a) (left) Contours of static pressure (upstream block) and water mass fractions (downstream

block). (b) (Right) The sonic surface colored by water mass fraction.

Flow Direction

Figure 6: Flow features of the DCC combustor near the combustion region. Purple regions indicate reverse flow; fuel injector plumes are indicated by fuel mass fraction iso-surfaces colored by temperature; streamlines denote recirculating flow in the flame-holding Base region. The Base region includes fuel injectors which are not visible.

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In Figure 7, we show the area averaged profiles of static pressure and Mach number obtained from the three-dimensional simulations and compare them to results from the one-dimensional model described above. The heat release profile, which is an input into the one-dimensional code, was independently estimated based on the injector locations and fueling parameters. The start of the shock train, which is an input to the isolator model, was assumed to be the leading edge of the isolator ramp. Other than these two assumptions, no attempt was made to fit the one-dimensional results to the three-dimensional averaged values. The figures show reasonable agreement between the three-dimensional and one-dimensional results everywhere except near the strut rear face. There is also a large discrepancy in the location of the downstream thermal throat. The reasons for this discrepancy are being investigated.

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Figure 7: Comparison of area-averaged three-dimensional results to one-dimensional results. (a) Static

pressure and (b) Mach Number

V. Discussion Two difficulties were encountered in the three-dimensional simulation. The first was isolator unstart; in particular, with fuel flow rates corresponding to an equivalence ratio of one, the shock train in the isolator moves all the way upstream to the inlet. For the preliminary results presented here, the fuel flow rates were reduced so that the shock train stayed downstream of the inlet near the leading edge of the isolator ramp. The full fuel flow rate condition is being studied to test the large backpressure margin of this combustor. The second difficulty was ignition. Since the inflow temperatures are fairly low, the fuel-air mixture does not auto-ignite. In the DCC test, several procedures are available to ignite the fuel-air mixture including the injection of Silane, and temporarily raising the inflow temperature. In the computations, ignition was facilitated by changing the combustion model, namely, lowering the activation energies of the various reactions in the kinetics model. Other methods, such as imposing a minimum threshold temperature are known to be less effective. Lowering the activation energies changes the ignition delay but does not change the net heat release of the reaction. Combustion modeling and simulation remains a challenge. Figure 8(a) shows the currently accepted12, 13, dominant reaction paths for laminar ethylene combustion in the temperature, pressure, and equivalence ratio

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ranges corresponding to ignition. Changing the reaction pressure and temperature will change the relative importance of these pathways and could easily introduce other reaction paths. Certainly, high temperature combustion or soot formation will involve dramatically different reaction pathways. The reaction pathways for JP-7 are more complicated yet; a current, practical reduced chemistry computational model for JP-7 contains 11 species and 14 reactions14. The simplified reaction mechanism used in the three-dimensional simulations—and shown in figure 8(b)—is valid at the conditions where it was calibrated, but cannot be expected to model the reaction mechanism over all conditions of engineering interest. Heat release is largely independent of this reaction mechanism; instead heat release is a function of reaction products and the gas model. However, ignition delay and flame holding are strongly dependent on this reaction mechanism; hence this simulation’s inability to predict flame holding.

C2H4 C2H5 M, H

Figure 8(a) Reaction path diagram for ethylene initiation and fuel consumption12, 13; (b) 6-species, 3-step reaction model used in the three-dimensional simulations.

Clearly, a reaction mechanism tailored to the combustor’s operating condition is desirable. Montgomery15 presents an automated optimization process for generating such reduced mechanisms. The previous discussion and results ignore the effects of turbulence on reaction rates. Other approaches need to be considered, such as Magnussen & Hjertager’s6 Eddy-Dissipation Concept (EDC), where the laminar reaction rates are discarded in favor of reaction rates proportional to the level of turbulence in the flow. The physical rationale is that turbulent mixing rates dominate the reaction process—instead of chemical kinetics. In summary, both the laminar kinetic approach and the EDC approach have their limitations. Combustion modeling continues to be an area of active research and debate with probability density functions (PDF) formulations2 leading the way.

VI. Summary Numerical simulations of varying levels of fidelity are being carried out in support of the NASA Glenn ISTAR DCC test. These include three-dimensional full Navier-Stokes solutions with finite-rate chemistry as well as one-dimensional Euler solutions with heat addition, area change, mass addition and friction. Preliminary comparisons between these models and with experimental data are encouraging. It is hoped that these numerical simulations, despite their uncertainties, will shed some light on the engineering uncertainties of the DCC test and reduce the risks of the program.

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Acknowledgments This work was supported by the NASA Computing and Information Communications Technology (CICT) program through the Computing and Interdisciplinary Systems Office (CISO) (contract NAS3-00145) at NASA Glenn Research Center. This work was also supported by the Engine Systems Technology Branch at Glenn Research Center (contract NAG3-2836). The authors would like to thank Jeffrey A. White of NASA Langley Research Center for help with VULCAN.

References 1. Lee, J., and Krivanek, T., “The NASA Integrated Systems Tests of an Airbreathing Rocket GRC-Direct Connect Combustor Experiment: Part 1 Design and Fabrication of the Experiment,” AIAA 43rd Aerospace Sciences Meeting, Reno, NV Jan 2005. 2. Baurle, R. A., “Modeling of High Speed Reacting Flows: Established Practices and Future Challenges,” AIAA 2004-0267, 42nd. Aerospace Sciences Meeting and Exhibit, Jan. 2004. 3. White, J. A., “VULCAN User Manual” Version 4.2.0, NASA Langley Research Center, March (2001). 4. Stewart, M. E. M., Suresh, A., Liou, M. S., Owen, A. K., Messitt, D. G., “Multidisciplinary Analysis of a Hypersonic Engine,” AIAA 2002-5127, AIAA (2002). 5. Mawid, M. A., Private Communication 6. Magnussen, B. F., and Hjertager, B. H., “On Mathematical Modeling of Turbulent Combustion with Special Emphasis on Soot Formation and Combustion,” 16th Symposium on Combustion, p. 719, (1976) 7. Suresh, A., Stewart, M. E. M., “Improving Airbreathing Engine Design Using Multidisciplinary Numerical Simulation,” JANNAF 27th. Airbreathing Propulsion Conference, Colorado Springs, (2003). 8. Sullins, G. and McLafferty, G., “Experimental Results of Shock Trains in Rectangular Ducts,” AIAA paper 92-5103, 1992.

9. Radhakrishnan, K., “LSENS-A General Chemical Kinetics and Sensitivity Analysis Code for Homogeneous Gas-Phase Reactions,” NASA RP 1328, January 1994.

10. Yungster, S. and Trefny, C.J., “Analysis of a New Rocket-Based Combined-Cycle Engine Concept at Low Speed,” AIAA paper 99-2393, 1999.

11. Oosthuizen, P. H., and Carscallen, W. E., “Compressible Fluid Flow,” McGraw Hill, (1997). 12. Varatharajan, B., and Williams, F. A., “Ethylene Ignition and Detonation Chemistry, Part 1: Detailed Modeling and Experimental Comparison,” Journal of Propulsion & Power, Vol. 18, No. 2, March 2002 13. Varatharajan, B., and Williams, F. A., “Ethylene Ignition and Detonation Chemistry, Part 2: Ignition Histories and Reduced Mechanisms,” Journal of Propulsion & Power, Vol. 18, No. 2, March 2002

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14. Staff, “Reduced Chemical Kinetics for JP-7,” First Report, Applied Physics Lab., The Johns Hopkins University, September 2004. 15. Montgomery, D. J., et. al., “Reduced Chemical Kinetic Mechanisms for Hydrocarbon Fuels,” Journal of Propulsion & Power, Vol. 18, No. 1, January 2002

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