EXPERIMENTAL STUDY OF A STRUT INJECTOR FOR CIRCULAR SCRAMJET COMBUSTORS

EXPERIMENTAL STUDY OF A STRUT INJECTOR FOR CIRCULAR SCRAMJET COMBUSTORS Christopher Rock Graduate Research Assistant and Joseph A. Schetz Advisor, Holder of the Fred D. Durham Chair Department of Aerospace and Ocean Engineering Virginia Tech, Blacksburg, VA, 24061 Supersonic combustion is a major challenge in scramjet engine design. Supersonic fuel injection and mixing research contributes to the effort to make the scramjet a viable option to power hypersonic aircraft, economical and reusable launch vehicles, and hypersonic missiles. An experimental study of a strut injector configuration was performed for application to high-machnumber scramjets with circular combustion chambers. The strut injector has sixteen inclined, round, sonic injectors distributed across four struts within a circular duct. The struts are slender, inclined at a low angle to minimize drag, and have two injectors on each side. The strut injector was experimentally studied under Mach 4, cold-flow conditions using two different molecular weight injectants, helium (molecular weight = 4) and air (molecular weight = 28.97). The primary goal of this study is the refinement of turbulence models for these complex mixing flows. Furthermore, injectant molecular weight has been identified as a parameter of critical importance in the development of the turbulence model upgrades. Experimental data such as presented here will be used to guide the continuing upgrade of turbulence modeling in a closely integrated program. Nomenclature A * = plume area of stoichiometric mixture C d = discharge coefficient d = jet diameter d eq = equivalent diameter G j = injectant mass flow rate M = Mach number P = static pressure P0 = stagnation pressure q = jet-to-free-stream momentum flux ratio R = resistance U = velocity w * = plume width I. Introduction In view of the very high freestream velocity of scramjets reaching Mach 10, fuel residence time is on the order of milliseconds 1 and supersonic combustion presents an interesting challenge in scramjet engines. It is, therefore, desirable to enhance penetration and mixing of the fuel plume in order to accomplish rapid combustion leading to a reduction of the required combustor length, reducing the skin-friction drag and heat transfer y = vertical distance from the duct centerline y * = vertical distance to injectant center of mass α = mass fraction of injectant ρ = density γ = ratio of specific heats Subscripts j = jet exit property = freestream property and increasing the net thrust. To improve the overall engine efficiency, the injection process must also induce low total pressure losses. Jet injector mixing enhancement in high-speed flows also has applications in other fields such as thermal protection systems and vehicle control by jet thrusters. Many injector configurations have been studied by various groups in an attempt to produce enhanced mixing and penetration. Some of these configurations can be seen in Figure 1 including Rock 1

wall jets, struts, and swept ramps. Extensive reviews of injector mixing characteristics are given in Schetz et al 2 and Kutschenreuter 3. Flush-walled injectors are often preferred over in-stream injectors because they minimize total pressure losses and heating, but some configurations can require the use of in-stream injectors in order to obtain adequate distribution of the fuel across the combustor. A circular combustor cross-section is one example where struts might be attractive. Of course, one has to reckon with the drag of the struts in assessing engine performance 3. Figure 1: Examples of various injector configurations (from Kutschenreuter 3 ) Very few of the detailed, high-speed mixing studies available in the literature concern injection and mixing in confined ducts representative of combustors, and one can expect that the effects of such confinement are very large. This is especially true for struts protruding into the flow. There are also bow shocks from the injection process itself. The purpose of the present research is to investigate the effectiveness of a four-strut injector configuration with multiple round, sonic injectors on each strut in a circular duct for application to high-mach-number scramjets with circular combustion chambers. The nominal Mach 4 air flow simulates conditions a scramjet combustor would encounter in Mach 10 flight. The general goals of cold-flow studies of injection and mixing in simulated scramjet combustors are first to determine if the penetration and mixing patterns observed are in agreement with those used for the injector design. Second, the experimental data can be used to gauge the uncertainty in computational predictions of such flows. The computational tools can then be used to design and analyze for hot-flow conditions with known uncertainty. The third and primary goal of this study is the refinement of turbulence models for these complex mixing flows. For this study, two different injectants were used, helium (molecular weight = 4) and air (molecular weight = 28.97), since injectant molecular weight has been identified as a parameter of critical importance in the development of the turbulence model upgrades. A. Test Facility II. Experimental Methods These experiments were conducted in the Virginia Tech blow-down type high-speed wind tunnel shown in Figure 2, which operates at speeds ranging from Mach 2 to 7. The blow-down type wind tunnel offers run times on the order of a few seconds at high Mach numbers with relatively steady flow conditions. This facility was obtained through our close and long-term collaborations with the Institute of Theoretical and Applied Mechanics of the Russian Academy of Sciences in Novosibirsk, Russia. Air (or other working gas) is supplied from a compressor to charge the storage bottles visible within the frame at the bottom. A special fast-acting control valve initiates flow into the plenum chamber. The flow then passes through Rock 2

a contoured, converging-diverging nozzle and out through the diffuser. Due to the working principle of the tunnel and the fast-acting control valve, there is a gradual decrease in total pressure during the run. The variation of the total pressure during the run is in the range of approximately 10%. For Mach numbers above 4, an electric heater raises the total temperature up to 800 K to prevent liquefaction. The nozzle exit diameter is 100 mm. The test cabin arrangement permits the use of relatively large in-stream models, especially at the higher Mach numbers. The wind tunnel setup for these experiments used a converging-diverging nozzle to achieve a nominal Mach 4 flow in the test section. Nominal flow conditions involve total pressure and temperature in the plenum chamber of 1317 kpa and 295 K. However, there is a weak oblique shock observed at the end of the nozzle, where the injector model attaches resulting in actual inflow conditions of Mach number, M = 3.9, total pressure 1311 kpa and total temperature 295 K. Figure 2: Layout of the Virginia Tech high-speed wind tunnel B. Injection System The injector setup investigated in this project resembles the combustion chamber of a scramjet engine. In a real scramjet, the combustor is situated downstream of the inlet and an isolator which compress the ingested air. In the experimental setup, the ducted strut injector model is mounted downstream of the convergent-divergent Laval nozzle of the high-speed wind tunnel. The injector model consists of a total of 16 injectors distributed over four struts within a circular duct and the necessary connections for the injectant supply. C. Ducted Strut Injector Model The ducted strut injector model is based on a circular duct extension of the tunnel nozzle. It contains four struts with 16 circular injection nozzles. Figure 3 shows a picture of the strut injector model. The struts have a width of 8.2 mm, they start at the front of the extension duct (i.e. the end of the tunnel nozzle), and they extend 148 mm in the flow direction with an inclination of 10. Two 1.52 mm circular nozzles on each lateral side of each strut create jets that penetrate into the tunnel crossflow at an angle of 30 relative to the streamwise axis of the duct. The centers of the injectors are located 92 mm from the leading edge Downstream of the injection position, a cylindrical flange connects the injector duct to the test cabin. A traversing system is installed on top of the test cabin, which positions flow measurement probes within the cabin. The injectant is supplied from a group of commercial gas bottles. The mass flow rate of the injectant is controlled using a system of two Teledyne-Hastings model HFC-D-307 digital mass flow controllers. Each mass flow controller uses a proportional integral derivative (PID) control valve. of each strut. The number, shape, and size of the struts were based on drag considerations and previous experience. The number, size, and location of the injectors were based on CFD studies. Due to the physical obstacle created by the struts, the formation of shocks at the edges and an expansion at the rearward facing edge of the struts can be expected. As a result, high total pressure losses in such a configuration are unavoidable. As pointed out before, an optimized injection system for a scramjet engine should combine good mixing efficiencies with low total pressure losses. A system including fixed flow obstacles has to compensate for this disadvantage by enabling Rock 3

better mixing in order to remain competitive with other geometries. A useful parameter for correlating transverse jet injection results is called the jet-to-freestream momentum flux ratio, q, defined as follows 2 ( ρu ) q 2 ( ρu ) j ( pγm = ( pγm 2 2 ) ) j (1) For this study, two experimental cases were run using different injectants. One case was run using helium injection, which safely simulates hydrogen fuel in a cold-flow, non-combusting environment. A second experimental case was then run using air injection. Each case was run with the same jet-to-freestream momentum flux ratio ( q ) to obtain a similar amount of fuel plume penetration. For the helium injection case, the total mass flow rate was set to G total = 22.5 g/s, which corresponds to q = 3.49 for this injector geometry and operating conditions. For the air injection case, the total mass flow rate was set to G total = 62.66 g/s to match the value of q = 3.49. These values of q are representative of good practice in strutted scramjet combustors. For both cases, the injectant jets are at sonic conditions and are highly underexpanded. Table 1 summarizes the injection parameters for the two experimental cases. Figure 3: Strut injector model Parameter Unit Helium injection case, G total = 22.5 g/s He G j [g/s] 1.41 3.92 q [-] 3.49 3.49 C d [-] 0.70 0.70 U j / U [-] 1.30 0.47 P0 j / P0 [-] 0.63 0.70 d [mm] 1.52 1.52 d eq = (C d ) 1/2 d [mm] 1.27 1.27 Air injection case, G total = 62.66 g/s air Table 1: Injection parameters for the two experimental cases (for a single injector) D. Concentration Sampling Probe In order to analyze the mixing of the helium with the air freestream, it is crucial to acquire accurate gas composition measurements. The concentration is measured in terms of the mass fraction of helium in the overall gas mixture. To determine this mass fraction, a special probe is used to simultaneously sample and analyze the gas mixture at a given position accurately. The fundamental concept of the gas analyzer used for this work was developed at Virginia Tech by Professor Ng 4. The concentration sampling probe is an aspirating type probe that is attached to a vacuum pump. A picture and diagram of the concentration probe are shown in Figure 4. The unit consists of a constant temperature hot-film sensor operating in a channel with a choked exit. The housing was designed to fit around the body of a TSI 1210-20 platinum sensor. The hot-film has a diameter of 50.8 µm and an active sensor length of 1.02 mm which is used in conjunction with a Dantec model 56C17 constant temperature anemometer (CTA) fitted with a Dantec model 56C01 CTA bridge. The overheat ratio of a hot-film sensor is defined as (R op - R 0 ) / R 0, where R op is the heated sensor resistance at operating temperature and R 0 is the cold sensor resistance at ambient temperature. An overheat ratio of 1.0 was used for the hot-film sensor of the concentration probe. The inlet hole at the tip of the probe has the same diameter as the choked orifice, 0.63 mm. These diameters were designed to preclude the occurrence of a standoff shock at the probe tip for supersonic flow due to the suction of the vacuum pump through the choked orifice. Schlieren flow visualization confirmed the absence of a standing normal shock. The internal probe diameter diverges from 0.63 mm at the inlet to 3.8 mm at the sensor plane, causing a normal shock to occur inside the probe in the diverging channel. By swallowing the shock into the internal diverging section of the probe, a stream tube equal in area to the probe capture area can enter the probe undisturbed and undistorted. Thus, an isokinetic sampling of the stream tube is accomplished. Through the Rock 4

diverging section and the normal shock inside the probe, the flow is decelerated to very low velocities. At Mach numbers around M = 0.05, the pressure and temperature measured inside the probe can be approximated to represent the total pressure and temperature of the fluid. The concentration sampling probe was calibrated to measure the helium molar fraction uniquely related to a given pressure, temperature and rate of heat transfer sensed at the hot-film. The hot-film responds to local mass flux variations. From the known helium molar fraction of the sample, the mass fraction can be calculated. The measurement uncertainty of the probe was found to be approximately +/- 0.005 for helium mass fraction measurements. Cone-Static Probe Concentration Probe Figure 4: Picture of the concentration sampling probe with an integrated cone-static probe (left) and diagram of the concentration probe (right) E. Cone-Static Pressure Probe The concentration probe has an integrated conestatic probe (see Figure 4). The cone-static probe is not required to determine the mixture composition, but the use of the cone-static probe allows for the determination of other quantities of interest for a given flow field using a multiple probe survey method. The cone-static probe was attached to the concentration probe to allow simultaneous concentration and cone-static measurements to be made. The cone-static probe consists of a 1.59 mm outer diameter pipe capped with a 10 half-angle cone. Four small pressure ports are located at 90 spacing around the surface of the cone. The conestatic probe is positioned in a location that is always outside of the oblique shock generated by the tip of the concentration probe. F. Miniature Five-Hole Probe A miniature, fast-response, conical, five-hole pressure probe is used to measure local values of Mach number, total pressure, and flow angularity. A picture and diagram of the five-hole probe are shown in Figure 5. The probe uses five miniature piezoresistive pressure transducers, which are located directly in its tip. The tip of the probe is a 45 half-angle cone with an outer diameter of 1.65 mm. Each pressure port has a diameter of 0.25 mm. The response time of the probe to a step input is about 11 ms. Two separate calibrations were performed to allow for the determination of Mach number and flow angularity. First, the five-hole probe was calibrated to determine the Mach Rock number of an airstream as a function of its port pressures. The calibration of the five-hole probe to determine Mach number is necessary due to the geometry of the probe. The probe has a blunt tip where the center (Pitot) port is located, which is surrounded by four peripheral ports. Downstream of the blunt tip, flow expansion occurs resulting in a region of lower pressure behind the tip in comparison to a sharp cone with the same halfangle. Beyond the region affected by flow expansion, the pressure distribution will quickly recover to that of a sharp cone. However, the peripheral ports for the probe used in the current study are located in the region affected by flow expansion. Sharp cone theory cannot be used to predict the readings for these ports and therefore, it is necessary to calibrate the probe to determine Mach number. The calibration was performed over a Mach number range of 1.6 3.9 and includes a total of 37 data points. This Mach number calibration is only valid for use of the probe in air, since mixture composition influences this calibration curve. The five-hole probe was also calibrated to determine the flow angularity of an airstream as a function of its port pressures. The calibration was performed at Mach 3.1 with an angularity range of +18 to -18 of pitch and 0-360 of roll. The angular calibration of the five-hole probe includes a total of 795 data points. The angular calibration is valid over a wide range of Mach numbers and it is also valid for use in both air and air-helium 5

mixtures according to the work of Swalley 5. Using a 40 half-angle cone in experiments run at a Mach number of 3.55 in air and 21 in helium, Swalley confirmed the theory that only one calibration curve is required to determine flow angularity over a wide Mach number range in either air or helium for this type of instrument. In addition, Centolanzi calibrated a 20 half-angle cone to determine the flow angularity of an airstream at Mach numbers of 1.72, 1.95, and 2.46 and also concluded that the effects of Mach number on the calibration map are either negligible or small 6. Figure 5: Picture of the five-hole probe (left) and drawing of the probe tip (right). All dimensions are in millimeters. G. Multiple Probe Survey Method and Data Analysis Procedure 1. Outside of the Region of the Fuel Jet Outside of the region of the fuel plume, the gas composition is known to be entirely air, so the concentration probe is not used in this region. The five-hole probe Mach number calibration is valid and the angular calibration is most accurate in this region. Therefore, the five-hole probe alone can be used outside of the fuel plume to determine local values of Mach number, total pressure, and flow angularity. The data reduction process needed to convert the port pressures into incoming flow properties follows that of Centolanzi 6. The measurement uncertainty of the five-hole probe was found to be approximately +/- 1.5% for Mach number, +/- 3% for total pressure, and +/- 1 for flow angularity in this region. 2. Within the Region of the Fuel Jet Inside the region of the fuel plume, the properties of the mixture must be accounted for and the data analysis procedure is more complex. First, a concentration probe survey is used to determine the local gas composition. The measurement uncertainty for the concentration probe is approximately +/- 0.005 for helium mass fraction measurements. Next, a method is needed to solve for the Mach number in the region of the fuel plume, since the five-hole probe Mach number calibration is not valid for gas mixtures. To determine the Mach number in this region, a multiple probe survey method is used. Corresponding data points are taken with the concentration probe, the cone-static probe, and the five-hole probe. The helium concentration data is then used with a combination of the Rayleigh-Pitot formula and a numerical solution of the Taylor- Maccoll equation for the local gas composition to determine the local Mach number at each measurement location. Once the Mach number at each measurement location is known, the total pressure and flow angularity can be solved for using the method of Centolanzi 6 as the five-hole probe flow angularity calibration is valid for airhelium mixtures. The measurement uncertainty is approximately +/- 2% for Mach number, +/- 5% for total pressure, and +/- 1 for flow angularity in the region of the fuel plume. III. Experimental Results A plane 178.3 mm (1.8 duct diameters) downstream of the circular injector centers was selected for data measurement purposes. At this measurement plane, which is about 2 mm beyond the end of the duct where the flow enters the test cabin, the flow field downstream of one half of one strut was surveyed. For the helium injection case, helium concentration, Mach number, and total pressure values were measured at the data measurement plane. For the air injection case, Mach number and total pressure values were measured. To check for symmetry, data points were also measured on the opposite side of the strut. The symmetry plane for the fuel plume was found to be shifted approximately 1 to 2 mm laterally relative to the centerline of the strut. This slight 1-2 mm shift of the fuel plume over a length of 178.3 mm is most likely attributed to a small, but undetected misalignment of the experimental hardware. Rock 6

A. Helium Concentration Results Results for the helium distribution presented as mass fraction contours across a section of the duct are shown in Figure 6. These mass fraction contours were determined from 127 experimental data points distributed across the fuel plume in the radial and peripheral directions. The projected outlines of the strut and the circular injectors are shown for reference. An examination of Figure 6 shows that at 1.8 duct diameters downstream, the injectant achieved good penetration across the combustor cross-section. However, the individual jets merged into a single large plume and the rate of mixing was somewhat slow. Calculations over the measurement grid provide parameters that characterize the plume and the mixing behavior, which are summarized in Table 2. Here, y* is the location of the center of mass of the injectant in the plume below the duct wall, w* is the maximum width of the equivalent stoichiometric hydrogen/air concentration contour, A* is the plume area within that contour, α max is the maximum injectant mass fraction, and y α,max is the distance from the duct wall to the location of α max. B. Mach Number and Total Pressure Results Figures 7 and 8 show contour plots of Mach number and total pressure at the measurement plane for the helium injection case compared to the air injection case. The projected outlines of the strut and the circular injectors are shown for reference. For the helium injection case, these plots were generated using the concentration data combined with 109 experimental data points taken with the cone-static probe and 186 data points taken with the five-hole probe according to the data analysis procedure described in Section II-G. For the air injection case, the contour plots were generated using 176 experimental data points taken with the five-hole probe. Figure 6: Contour plot of helium concentration at a plane 1.8 duct diameters downstream of the circular injector centers Parameter Unit Value y * [mm] 16.9 w * [mm] 24 A* [mm 2 ] 408 α max [-] 0.101 y α,max [mm] 16 Table 2: Strut injector mixing parameters (for an entire plume created by 1 strut with 4 injectors) Rock 7

Helium Injection Air Injection Figure 7: Contour plots of Mach number for helium injection vs. air injection at a plane 1.8 duct diameters downstream of the circular injector centers Helium Injection Air Injection Figure 8: Contour plots of total pressure for helium injection vs. air injection at a plane 1.8 duct diameters downstream of the circular injector centers A complex shock system forms downstream of the injector array, which includes oblique shocks from the struts and bow shocks from the injectors. This phenomenon reduces the Mach number of the flow downstream of the strut as shown in Figure 7. Also, the injectant jets are at sonic conditions, whereas the freestream is nominally at Mach 4 conditions. The region of reduced Mach number in the vicinity of the fuel plume is substantially larger for the helium injection case compared to the air injection case. This is due to the properties of the gases such as molecular weight and specific heat ratio. These properties influence the sound speed of a gas, therefore helium has a higher sound speed than air at the same temperature conditions. Mach number is inversely proportional to sound speed, so the Mach number in the region of the fuel plume is lower for the helium injection case. Both cases showed good penetration of the injectant across the combustor cross section, but there is a substantial total pressure loss downstream of the strut as shown in Figure 8. The total pressure loss is larger for the helium injection case than for the air injection case, which is largely due to Mach number effects. Another factor that contributes to the lower total pressure for the helium injection case is the total pressure in the injector manifold, Rock 8

which is lower for the helium injection case than for the air injection case (see Table 1). IV. Application of Experimental Research to CFD Turbulence Model Upgrades The experiments presented here are part of an integrated experimental and computational study being conducted by a team of researchers from Virginia Tech and a small business, CRAFT Tech (Combustion Research and Flow Technology, Inc.). The experimental research is being conducted at Virginia Tech, whereas the computational research is being conducted at CRAFT Tech. The primary goal of this study is to upgrade the turbulence models that are used for CFD predictions of the flow inside a scramjet combustor. There are two primary upgrades that are currently being developed for the k-ε turbulence model: (1) scalar fluctuation modeling and (2) a baroclinic turbulence source term correction. Scalar fluctuation modeling predicts local variations in turbulent Prandtl and Schmidt numbers. CFD predictions for the flow inside a scramjet combustor currently generally use turbulence models that utilize global estimates for the turbulent Prandtl and Schmidt numbers. In this type of flow, the local values of turbulent Prandtl and Schmidt numbers can vary significantly and this is believed to be a source of error. The baroclinic torque term is a modification to the turbulence model to account for the effects of strong density gradients, which occur in high-speed mixing flows. In high-speed mixing flows, the flow conditions deviate significantly from those used in the standard k-ε model derivation. Therefore, refinement of the turbulence model is necessary to accurately predict these complex mixing flows. To support the turbulence model upgrades, experiments are required that parametrically vary the jet molecular weight and freestream Mach number, while keeping other basic mixing parameters the same. The experiments provide a database of fuel injection and mixing data that is being used for turbulence model refinement and CFD code validation. Figure 9 shows an example of how the experimental data is being used to upgrade the turbulence modeling. In this figure, the experimental helium concentration results along the centerline of the strut injector are compared to CFD predictions with and without the additional baroclinic torque term. The baroclinic torque term was tested and validated for a wide range of experimental cases consisting of different flushwall and in-stream injector designs, varying molecular weight fuels, and varying freestream Mach numbers 7. From examining Figure 9, it is evident that the baroclinic torque term significantly improves the CFD predictions. Figure 9: Comparison of the mixing results predicted by the CRUNCH CFD code to the experimental data along the centerline of the strut for the helium injection case V. Conclusion This paper presented the results of an experimental study of a four-strut injector configuration with multiple round, sonic nozzles on each strut in a 100 mm diameter circular duct under cold-flow conditions for application to high-mach-number scramjets with circular combustion chambers. The freestream Mach number was nominally 4, which simulated the conditions a scramjet combustor would encounter in nominal Mach 10 flight. The primary goal of this study is the refinement of turbulence models for these complex mixing flows. Injectant molecular weight has been identified as a parameter of critical importance in the development of the turbulence model upgrades and the use of two different injectants, helium and air, Rock 9

allowed the effects of injectant molecular weight to be studied. For comparison purposes, a constant jet-to-freestream momentum flux ratio was maintained to achieve a similar amount of fuel plume penetration for the two experimental cases. The main reason for considering an intrusive injector design for application in such a challenging thermal environment as a Mach 10 scramjet combustor is the goal of minimizing the combustor length in the low aspect ratio combustors currently being considered for such applications. An in-stream injector inherently yields better penetration and airstream coverage in short axial distances than a flush-wall injector. If adequate mixing can also be achieved in a short distance, then the drag and thermal load penalties of the in-stream injector can be overcome. For the case with helium injection, the use of helium safely simulates hydrogen fuel in a noncombusting environment. The experimental results for the helium injection case obtained at approximately 1.8 duct diameters downstream showed good penetration of the injectant across the combustor cross-section, but the individual jets had merged into a single large plume and the rate of mixing was somewhat slow. In addition, a substantial total pressure loss occurred in the flow downstream of the strut. There were substantial regions of fuel-rich (using the stoichiometric ratio of hydrogen to air, 0.0292, as a metric) concentrations in the plumes, even though the overall helium-air mixture in the duct was lean, on the same basis. The maximum concentration of helium detected by the experiment was 10.1%. One might have expected better mixing based on simple, isolated injector correlations. The ratio of the measured injectant mass flow to the calculated air flow (equivalent fuel/air ratio) in the duct was References 0.0098 for the helium injection case, which corresponds to an equivalence ratio of 0.34 in terms of hydrogen. The area of the helium plume created by one strut with four circular injectors was calculated to be 408 mm 2 based on the experimental data. Thus, the overall helium plume area in the duct covered by a total of 16 jets distributed across four struts is 1632 mm 2, which accounts for 20.8% of the total duct cross section. The air injection case exhibited similar features to the helium injection case including good penetration of the injectant across the combustor cross-section, the individual jets merging into a single large plume, and substantial total pressure loss in the flow downstream of the strut. However, the helium injection case had a larger overall total pressure loss than the air injection case. The equivalent fuel/air ratio for the air injection case was 0.0273 in comparison to 0.0098 for the helium injection case. The increase in the fuel/air ratio for the helium injection case vs. the air injection case is similar to the fuel/air ratio increase that would occur for a scramjet engine operating on a low molecular weight fuel vs. a high molecular weight fuel. The experiments presented here are part of an integrated experimental and computational study that is being conducted to improve the turbulence models that are used for CFD predictions of the flow inside a scramjet combustor. A database of fuel injection and mixing data is being built for turbulence model refinement and CFD code validation. Certain improvements to the turbulence modeling have already been achieved as part of this integrated experimental and computational study. Nevertheless, it will likely be possible to further improve the turbulence modeling by obtaining additional experimental data. 1. Maddalena, L., Campioli, T.L., Schetz, J.A., Experimental and Computational Investigation of an Aeroramp Injector in a Mach Four Cross Flow, AIAA/CIRA 13 th International Space Planes and Hypersonics Systems and Technologies, AIAA 2005-3235, June 2005. 2. Schetz, J.A., Thomas, R.H., and Billig, F.S., Mixing of Transverse Jets and Wall Jets in Supersonic Flow, IUTAM Symposium on Separated Flows and Jets, Novosibirsk, July 1990. 3. Kutschenreuter, P., Supersonic Flow Combustors, in Scramjet Propulsion, (E.T. Curran and S.N.B. Murthy, Editors), AIAA, New York, 2000. 4. Ng, W.F., Kwok, F.T., and Ninnemann, T.A., A Concentration Probe for the Study of Mixing in Supersonic Shear Flows, AIAA paper 89-2459, July 1989. 5. Swalley, F.E., Measurement of Flow Angularity at Supersonic and Hypersonic Speeds with the Use of a Conical Probe, NACA TN D-959, September 1961. 6. Centolanzi, F.J., Characteristics of a 40 degree Cone for Measuring Mach Number, Total Pressure, and Flow Angles at Supersonic Speeds, NACA TN 3967, May 1957. 7. Ungewitter, R.J., Brinckman, K., and Dash, S.M., Advanced Modeling of New Fuel/Air Mixing Data Sets for Scramjet Applications, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Denver, CO, AIAA-2009-4940, August 2009. Rock 10