Roland H. Krauss

Experimental Study of a Dual-Mode Scramjet Isolator

A constant-area isolator was fabricated and tested in conjunction with a Mach 2 hydrogen-air combustor operating at a simulated Mach 5 flight enthalpy. Predicted isolator performance was validated through pressure measurements obtained via low-frequency pressure taps. The maximum pressure ratio measured in the combustor approached the design limit of 4.5. Scramjet operability, the range of equivalence ratios over which combustion was sustained without shock-inlet interaction, was improved to 0.06-0.32, as opposed to 0.32-0.37 without the isolator. For a given change in fuel equivalence ratio, the location of the shock train was easier to control with the isolator modification. Shock-train location repeatability was found to vary somewhat with equivalence ratio. Small fluctuations in the time-resolved pressure history indicated that the shock train was relatively temporally steady for a given equivalence ratio. High-frequency pressure measurements were within a 95% confidence interval of low-frequency pressure measurements. High-frequency results indicated that an increase in pressure and large pressure fluctuations occurred near the leading edge of the shock train. Power spectral analyses also indicated that there is significant variation in the frequency content of the pressure signal upstream and downstream of the shock-train leading edge. These results suggest that methods of shock-train leading-edge detection may be developed using pressure-time history characteristics other than the pressure magnitude.

Experimental and Numerical Study of a Dual-mode Scramjet Combustor

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

Shock Train Leading-Edge Detection in a Dual-Mode Scramjet

Time-resolved pressure measurements in a dual-mode scramjet isolator were examined to investigate the potential for using the measurements for shock train leading-edge detection. Changes in the pressure magnitude, standard deviation levels, and frequency content were observed as the shock train advanced upstream past each pressure measurement station. Three detection criteria were defined and examined: 1) 150% of the normalized pressure magnitude upstream of combustion influences, 2) 150% of the normalized pressure standard deviation level upstream of combustion influences, and 3) the maximum value of the normalized pressure standard deviation. Another method of shock train leading-edge detection involved the examination of the frequency content of the pressure signal using power spectra analysis. Results indicated that the second detection criterion provided the earliest method of shock train detection as the shock train moved upstream, followed by the first and third criteria. Also, the frequency content of the pressure signals significantly changed near the shock train leading edge. However, a comparison of this method to the three criteria first examined showed that it did not provide earlier shock train detection.

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

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.

Test gas vitiation effects in a dual-mode scramjet combustor

An experimental study was conducted to characterize the influence of combustion air preheater major vitiate species (H 2 O and CO 2 ) on scramjet combustion. These species were added to an initially clean airflow that was supplied by an electrically heated facility. With dry air, the scramjet combustor operated in the supersonic mode at an equivalence ratio in the range of 0.25-0.32 and transitioned to dual mode over an equivalence ratio range of 0.35-0.37. At an equivalence ratio of 0.27, the combustor operated in the supersonic mode for three cases: 1) dry air, 2) air vitiated with 5% H 2 O by mole, and 3) air vitiated with 5% H 2 O and 2.5% CO 2 by mole. In the second case, the combustor pressure distribution decreased 10% relative to dry air and, in the third case, another 2% decrease was measured. At an equivalence ratio of 0.35, the combustor operated in the dual mode with dry air, but in the supersonic mode with 7 % H 2 O. This is the first demonstration of mode transition solely caused by test gas vitiation. It is therefore important to account for such effects when extrapolating from vitiated ground testing to flight.

Laser Selection Criteria for OH Fluorescence Measurements in Supersonic Combustion Test Facilities

The limitations of the application of dye laser and narrow-band tunable KrF excimer laser systems to OH planar laser-induced fluorescence measurements in supersonic combustion test facilities are examined. Included in the analysis are effects of signal strength, collisional quenching, beam absorption, and fluorescence trapping on achievable measurement accuracy using several excitation and detection options for either of the two laser systems. Dye based laser systems are found to be the method of choice for planar imaging when the line integral of OH concentration along the incidence or detection paths is less than 10 16 cm/cm 3 , whereas the ArF based systems provide significant reduction in measurement ambiguity when either concentration line integral is in excess of 10 16 cm/cm 3

Mixing and Chemical Kinetics Interactions in a Mach 2 Reacting Flow

An experimental study of transverse hydrogen injection and combustion behind a rearward-facing step into a Mach 2 airflow was conducted in an electrically heated (not vitiated), continuous-flow facility to evaluate the effects of initial conditions (temperature, pressure, and equivalence ratio), and analyze the interactions between mixing and combustion in supersonic, reacting flows. Neither mixing nor reaction rates dominate in this particular regime, thus the use of initial conditions to scale fuel mixing (i.e., dynamic pressure ratio) has to be modified by a descriptor that includes the effects of combustion on the flow conditions at the fuel injection station. Combustor inlet static pressure was varied from 0.25 to 0.5 atm, and total temperature from 300 to 850 K. Injector configurations include both single and staged injection with injectors of 1 and 1.5 mm diam, transverse to the airflow, behind a 5-mm rearward-facing step. Images of visible flame emission distribution at several temperatures correlated with pressure and temperature measurements are used to describe the coupling between fluid dynamics and chemical kinetics, discussed in terms of a characteristic global Damkohler number (ratio of chemical reaction rate to turbulent mixing rate). A proposed modification to mixing scaling with dynamic pressure ratio in the presence of heat release effects is presented.

Planar OH density and apparent temperature measurements in a supersonic combusting flow

Distributions of the absolute density of OH and of the population densities of five OH rotational states were obtained in a supersonic H 2 -air combustion tunnel using planar laser-induced fluorescence. A tunable KrF excimer laser was used to excite, nonsimultaneously, the five separate ro-vibronic lines in the (3-0) vibrational band of OH. Measurements were calibrated against measurements in an atmospheric air furnace, where known quantities of OH are formed by the thermal dissociation of H 2 O molecules. Despite flame intermittency, a sum of each of these nonsimultaneous measurements represents the time-averaged population density of the probed state. Present results show that these time-averaged measurements of the populations of the probed states still follow a Boltzmann distribution. Therefore, measurements of the average population of any pair of states can provide the two independent average properties required to define the average thermodynamic state of OH.

Calibration source for OH laser-induced fluorescence-density measurements with thermally dissociated H2O in atmospheric air.

A calibration technique for OH laser-induced fluorescence (LIF) density measurements through the use of the thermal dissociation of ambient H(2)O in an atmospheric air furnace with a tunable KrF laser has been demonstrated. The stable and uniform concentration of OH produced in the furnace permits direct calibration of LIF signals without the uncertainties associated with reference flames. The presence of OH in atmospheric air that is heated to temperatures exceeding 1500 K is sufficient for LIF measurements with most OH LIF laser systems. The measured OH density is found to agree well with the computed OH chemical-equilibrium density over a temperature range of 1500-1850 K.

Surface temperature-field imaging with laser-induced thermographic phosphorescence

A technique to remotely image temperature distributions of heated metallic surfaces is extended to higher temperatures. It uses a Dy(+3):YAG thermographic phosphor (TP) bonded to the surface and excited by radiation at 355 nm. Digital images of the emission from two excited states were recorded and divided by each other to correct by normalization for illumination and coating nonuniformities. Results show that the TP can survive heating and cooling cycles to 1400 K and that emitting states achieve thermodynamic equilibrium before radiating. Temperatures in the range of 300-1300 K were determined by normalization of pairs of emission images with a single calibration constant. Uncertainties of +/-7-13% at a spatial resolution of 20 microm and +/-0.7-4% at a resolution of 500 microm were achieved.