Tetsuo Hiraiwa

Mach 6 Testing of a Scramjet Engine Model

Testing of a subscale scramjet research engine model was carried out in the Mach 6 Ramjet Engine Test Facility of the National Aerospace Laboratory, Kakuda Research Center. The engine had a sidewall compression-type inlet. The fuel was hydrogen. With the attachment of a short strut on the top wall, intensive combustion with high-combustion efe ciency was attained, and the engine-produced thrust canceled the drag. The e ame was held in the low-velocity region around the step, even after the ignitors had been turned off. When the fuel e ow rate was small, there was a different combustion mode with weak combustion and little thrust. The unstart condition seemed to begin around the cowl. Tangential injection of fuel inhibited intensive combustion.

Analyses and Application of Gas Sampling to Scramjet Engine Testing

Gas sampling has been used in combustor studies and in scramjet engine testing. Because the gas sampling is based on the assumption that the gas composition is frozen in the sampling process, the critical Damkohler numbers necessary to quench reactions in the gas-sampling probes were evaluated using a reduced kinetic model. The phase plane analysis showed that reactions in probes can be extinguished if the probe Damkohler number is less than about 10. The analytical results were cone rmed by numerical calculations using full kinetics. The shock swallowing into sampling probes was examined using numerical simulations for the low-Reynolds-number e ow. These theoretical results were verie ed by experiments using four kinds of probes with various cone gurations in a Mach 2.5 supersonic combustor. Based on the results, e ne sampling probes with a tip diameter less than 0.3 mm are recommended for scramjet testing. Based on these calibration studies, gas sampling was successfully applied to scramjet engine testing under a e ight Mach number up to 8, to reveal interesting features in the internal e ow in swept-back engines.

Drags in Scramjet Engine Testing: Experimental and Computational Fluid Dynamics Studies

No-fuel internal drag of a side-compression scramjet engine was evaluated by using two one-e fth-subscaled models, one forwall pressure measurement and theother forforce measurement under conditionsof a e ight Mach number of 4. The pressure and frictional drags in various parts of the models were estimated from these windtunnel tests. Comparison between the pressure measurement and the force measurement revealed that the drag derived by these wind-tunnel tests agreed within 5%. After examining the consistency between the pressure and the force experiments, these results were used to calibrate a newly developed computational e uid dynamics code. The frictional drag and the heating rate on the engine internal walls were evaluated with the unstructured-grid code to be compared with those obtained from the one-e fth-subscale model and the full-scale engine. The total drag coefe cient of the scramjet engine, including the installation drag, was found to be 0.281 and the internal drag coefe cient was found to be 0.093. Consequently, two-thirds of the total drag measured in engine testing in the Ramjet Engine Test Facility was produced by the external e ow over the engine module. Subtracting the external drag, the internal performance delivered by the H 2-fueled scramjet engine is discussed.

Vitiation Effects on Scramjet Engine Performance in Mach 6 Flight Conditions

Quasi-one-dimensional analyses with either chemical equilibrium or finite-rate reaction were conducted to evaluate the effects of contamination due to flow vitiation on engine performances, namely, thrust production. The incoming flow state was calculated based on measurements, and a higher total enthalpy for the freestream through a vitiation heater compared to that through a storage heater was found, the difference in the flow condition being responsible for two-thirds of the reduction in the thrust production with vitiation observed in the engine tests. With the finite-rate reaction calculation, the possibility of a further reduction due to the contamination effect on the combustion mechanism was found, which was responsible for one-third of the measured reduction. The one-dimensional analyses were further pursued to find matched test conditions with the storage heater to that with the vitiation heater in view of the thrust production and pressure distribution within the engine, and both freestream enthalpy and fuel equivalence ratio should be adjusted to attain the matched conditions.

AUTO IGNIT-ION IN A SUPERSONIC COMBUSTOR WITH PERPENDICULAR INJECTION BEHIND BACKWARD-FACING STEP

Flow field at pre-ignition phase in a supersonic combustor with perpendicular injections behind a backward facing step was simulated in cold flow tests with inert gas injection, to investigate the mechanism of the ignition enhancement observed in the ignition tests with the same combustor. The interaction, namely the merging, of separation at step base with that upstream of the injector caused enlargement of ignition region and the enhancement. The change in ignition ability was estimated based on ignition parameters. Based on the experimentally obtained ignition parameters, the ignition limit for the interacted separation region was predicted and showed reasonable agreement with the experimental results.

Detonation engine development for reaction control systems of a spacecraft

Study of a Rotating Detonation Engine (RDE) has been carried out in many research institutions. The RDE has the advantage of high efficiency and simple structure. Taking this advantage of its benefits, application to the reaction control system of a spacecraft is expected. Toward the practical use, the knowledge of the detonation propagation in operation and quantification of stable operating range is essential. In the present study, we have produced the RDE with the linear channel that specializes in visualization. The first experimental results of this RDE are reported. The main purpose is obtaining basic combustion characteristics. The optical access of the RDE can be easy by linear channel. In the experiment, the propagation of the detonation wave front is confirmed by a high speed video camera and pressure sensors. The location information of the wave front obtained from the image, the x-t diagram after ignition was obtained. Thereby, the transition to detonation is confirmed. The propagation of the multi-wave front in the steady state was confirmed by the images and pressure sensor. Propagation velocity of the wave front is also obtained by the pressure history and the images.

Scramjet inlet flow computations by hybrid grid method

Computations of internal viscous flowfields of scramjet models were conducted at inflow Mach number of 5.4. An unstructured hybrid grid method was used to compute complex geometries such as scramjet models with a short strut. The numerical method to solve the Navier-Stokes equations on the hybrid grid was developed using a finite volume cell vertex scheme and the LU-SGS implicit time integration algorithm. The computational results using one-equation turbulence models showed good agreement with the experimental data. The flow features and the changes of flowfields due to the short strut located in the upper passage were discussed. It was revealed that a thick subsonic region did not exist in the combustor near the top wall at Mach number 5.4. It was favorable features to avoid the engine unstart. With the strut, relatively low velocity regions became larger and the down wash flow toward the cowl behind the step became strong. The overconcentration of the fuel toward the top wall during the weak combustion was found in the experiment. From the computational results, the reason of this overconcentration was realized that the airflow near the injector was turned to the top wall due to the small influence by the combustion in the experiment. The computational time and the accuracy of the present method were the same level as the conventional structured grid methods. Thus the present method seemed to be engineeringly very useful for analysis and design of the high-speed propulsion engines.

Geometrical effects to aerodynamic performance of scramjet engine

A side wall compression type scramjet model was tested at Mach 4, 6 and 8 flight conditions. The effects of the incoming Mach number on the aerodynamic performance were investigated with a simple model, and it was found that the compression ratio was not high enough in M6 and M8 cases. Additional structures like a strut were applied to the model to enhance the pressure ratio. The results showed that it would achieve higher pressure in some regions, but always associated with the local pressure decrease. The aerodynamic coefficients study of every configuration tested here showed the modifications with these structures would increase the drag coefficient by the factor of up to 1.3. The modification should be made with a proper configuration , at proper location, considering with the combustion requirement.

EXPERIMENTS ON A SCRAMJET ENGINE WITH RAMP-COMPRESSION INLET AT MACH 8 CONDITION

National Aerospace Laboratory of Japan (NAL) has been testing sidewall-compression type scramjet engines under Mach 8 flight condition since 1995. Over 100 tests under the condition have been conducted at NALs RamJet- engine Test Facility (RJTF). However, no significant achievement has been taken from these experiments. Several type/configuration of engines were delivered and tested, but they could not produce thrust larger than their internal drag. This report presents recent modification to one of the engines and its experimental achievements: To reduce the drag caused by a strut, which has been generally used as a flame holder and a part of air compressing system, a ramp-compression wall (ramp block) is installed on the topwall instead of the strut. Then, this engines fundamental performance and characteristics are displayed and discussed. Additionally, we show some experimental results of heat flux on the ramp. Comparing these results with our former experimental results, we discuss advantages of the ramp block model.

Optical Measurements of Ethanol/Liquid Oxygen Rocket Engine Combustor with Planar Pintle Injector

The pintle injector is a promising candidate of the propellant injection systems for a rocket engine with deep throttling capability which is essential for future space transportation missions. However, studies focusing on combustion phenomena in a rocket engine with a pintle injector is rather limited. In this study, optical measurements inside an ethanol/liquid oxygen rocket engine combustor with a planar pintle injector are conducted to clarify spray and flame structures in the pintle injector. Combustion tests where the chamber pressure is 0.36 to 0.40MPa and O/F is 1.15 to 1.40 are conducted. Effects of the injection configuration on spray structures are also evaluated. High speed imaging techniques are used to observe the flame and spray structures under hot fire conditions. Strong chemiluminescence of CH is observed in the vicinity of the impinging point of two propellants. Luminous flame is observed intermittently in the vicinity of the faceplate and the upper wall of the combustor with the direct observation of the flame. Characteristic exhaust velocity efficiency with oxidizercentered configuration is lower than fuel-centered configuration due to the large amount of propellant impinging on the upper wall of the combustor. Periodical phenomena with the frequency of approximately 300Hz which can be related to atomization processes are also observed.