Search

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2019-01-03, 資料番号: PA1910059000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2018-04-16, 資料番号: PA1810060000

The pintle injector is a promising candidate for propellant-injection systems for various rocket engines. However, fundamental studies focusing on the pintle injector are rather limited, and the effects of key design parameters on combustion behaviors of the injector are still not identified. Therefore, combustion tests of an ethanol/liquid-oxygen rocket-engine combustor with a planar pintle injector are conducted to investigate the effects of total momentum ratio and O/F on the combustion characteristics of the pintle injector. The total momentum ratio and O/F vary from 1.0 to 2.2 and from 0.32 to 3.65, respectively, whereas the combustion pressure is kept constant at approximately 0.4 MPa. The spray structures are observed with high-speed cameras. The characteristic exhaust velocity (C*) efficiency increases with the increase in O/F in the fuel-centered configuration, and the opposite tendency is observed in the oxidizer-centered configuration. The C* efficiency decreases with the increase in the total ...

This paper describes a concept study of an optional kick stage system named “PLUS (for Planetary mission, Long-duration small-thrust Upper Stage)” equipped with a relatively small-thrust engine for interplanetary missions. The thrust force of PLUS assumed in this study is a maximum 29.4 kN, which is relatively small, used to inject a payload into interplanetary orbit with sufficiently low gravity loss i.e. less than 2%, although two split delta-Vs such as the Russian Proton/Breeze-M will improve performance efficiency (Proton Launch System Mission Planner's Guide, 2009 [1]). The orbital maneuvering capability of PLUS system is evaluated through Mars orbital maneuvering simulations. In this study, two types of PLUS systems are assumed. First, a 29.4 kN thrust force kicks stage system designated as “PLUS1” to inject the 3000 kg main payload into Mars transfer orbit. Second, a 9.8 kN thrust force small PLUS system designated as “PLUS2” to inject a 500 kg secondary payload into Mars transfer orbit. In the first case, the above-mentioned split delta-Vs are conducted to inject the main payload into Mars transfer orbit with sufficiently low gravity loss. In the second case, on the other hand, PLUS2 attached to the secondary payload is dual launched together with a primary payload into a geostationary transfer orbit, GTO, whereupon PLUS2 is initiated slightly before perigee to inject the secondary payload into Mars transfer orbit utilizing Electric delta-V Earth Gravity Assist, EDVEGA scheme via the on-board Ion Electric propulsion System, IES (Kawaguchi, 2001 2 and 3). Throughout the simulations, some optimized configurations of PLUS system covering a wide variety of space missions are suggested in this paper., 資料番号: PA1410024000

Aero-vibro acoustic simulation for the prediction of harmful acoustic loading at lift-off of launch vehicle is developed. In this simulation technique, high-fidelity large-eddy simulation with computational aeroacoutics based on full-Euler equations is employed for computing jet aeroacoustics and their propagation to the outside of payload fairing. Acoustic field inside the payload fairing is computed by the coupled vibro-acoustic simulation based on finite element method. A simplified fairing model is used for the validation of the present method. An impact hammer test and acoustic vibration test using a loudspeaker in an anechoic chamber are conducted for validating the structural model. Then, the accuracy of this method is validated by using the acoustic vibration test result with a subscale rocket engine.

資料番号: PA0810079000

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.

著者人数: 4名, 資料番号: AC1000033000

著者人数: 5名, 資料番号: AC1000030000

A one-dimensional theoretical method for an evaluation of the net thrust of a scramjet engine is presented. To calculate the thrust performance of the engine analytically, an efficiency characterizing compression process, such as recovery factor in total pressure or kinetic energy efficiency etc., is necessary. In the previous theoretical method, this efficiency was roughly assumed. Although the effect of aerodynamic parameters and configuration parameters on the thrust can be investigated with this rough assumption, the net thrust evaluated with this method was not enough accurate to be compared with the measured thrust in experiments. The total pressure loss is chosen as the efficiency in this paper. This total pressure loss is evaluated from the engine internal drag with the wind tunnel experiment, which is due to irreversible process through the engines. The theoretical maximum thrust of the engines can be calculated with this new method.

A modified sidewall-compression-type scramjet engine was tested under M8 flight condition. A strut with a long tail led to a 60% higher maximum thrust increment than a strut with a truncated tail, while allowable fuel flow rate (i.e., the flow rate at engine unstart limit) was almost identical. To increase the allowable fuel flow rate and the maximum thrust increment due to combustion without engine unstart, distributed injection in vertical direction and installation of ramps in the isolator between the inlet and the combustor were applied. Distributed fuel injection led to a 30% higher maximum thrust increment than the uniform fuel injection, and the installation of the ramps also gave a 30% higher one than that in the engine without ramp. Gas-sampling at the exit of the engine with distributed fuel injection showed 88% of bulk combustion efficiency. Specific impulses of the maximum thrust increments, δIsp , of the modified engine were 60–80% higher than that of the original engine.

51st AIAA/SAE/ASEE Joint Propulsion Conference (July 27-29, 2015.Hilton Orlando) Orlando, Florida, USA., 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., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AC1600102000, レポート番号: AIAA 2015-3845

To complement our hydrogen-fueled scramjet engine research, aerodynamic tests were conducted by using one-fifth-subscaled (0.44 m-long) models in Mach 6.7 freestream conditions. The drag of the engines was estimated by using a force balance and pressure distributions on the internal wall to separate the internal drag and the external drag. The swept shock wave theory yielded the air capture ratio and the additive drag due'to the spilled flow. Two methods for drag evaluation, based on force balance measurement and on wall pressure Pw measurement, were compared to examine their accuracies. They enable identifying the origins of pressure and friction drags. It was found that frictional drag constituted more than half of the internal drag of engines and that the pressure recovery across the engine decreased to about 4% when large struts were installed in the inlets. The fundamental principles and the experimental validation of the methods for evaluating the internal drag in scramjet engines with large struts is reported.

To predict harmful acoustic loading observed at lift-off of the launch vehicle, aero-vibro acoustic simulation technique is developed. High-fidelity large-eddy simulation with computational aeroacoustics based on the full Euler equations in time domain are employed to predict generation of the acoustic waves and their propagation to the fairing. Coupled vibro-acoustic analysis based on finite element method are applied to compute the transmitted acoustic wave into the fairing both in the time and frequency domains. Acoustic measurement of a simplified fairing model with a subscale liquid rocket engine is conducted, and validation study of the present technique is performed. Reasonable agreement is obtained for peaks of the acoustic spectrum taken inside the fairing model. Such peaks are found to be related to the internal acoustic modes and ring modes of the faring structure. However, further study is required to obtain quantitative agreement.

A subscale acoustic test, the H3-scaled Acoustic Reduction Experiments (HARE), was conducted to predict liftoff acoustic environments of the H3 launch vehicle currently being developed in Japan. The HARE is based on 2.5% scale H3 vehicle models, which is composed with a GOX/GH2 engine and solid rocket motors, Movable Launcher (ML) models with upper deck water injection system and Launch Pad (LP) models with deflector and lower deck water injection systems. Approximately 20 instruments measured far/near field acoustic and pressure data. Preliminary results are presented in this presentation.

Computations of internal viscous flowfields of scramjet models were conducted at inflow Mach numbers of 3.4 and 5.45. A hybrid unstructured grid method was used to compute the scramjet models with and without a short strut. The numerical method to solve the Navier-Stokes equations on the hybrid unstructured grid was developed using a finite volume cell vertex scheme and the lower-upper symmetric Gauss-Seidel (LU-SGS) implicit time integration algorithm. The computational results revealed that a thick subsonic region did not exist in the combustor near the top wall at Mach number 5.45. It was a desirable feature to avoid unstarting the engine. With the use of a strut, the relatively low velocity regions increased and the downward flow toward the cowl behind the step became strong. An overconcentration of fuel toward the top wall during weak combustion was observed in the experiment. The reason for this was that the airflow near the injector was turned to the top wall due to the small influence of combustion in the experiment.

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.

著者人数: 5名, 資料番号: AC1000021000

We conducted a sensitivity analysis of engine performance on measurements of gas sampling, pitot pressure and static pressure. It indicates that the heat loss in engine and the insufficient quenching of reactions in the gas sampling probes results in the mismatch in the mass flux through the engines. An analytical study of the reactive flow in probes showed that the introduction of a small heat of reaction with radical recombination to the energy equation yielded a saddle-type singularity in a phase plane for temperature and radical. The separatrix running from the initial point to the saddle point gave the criterion dividing the thermal runaway and the chemical quenching. A one-dimensional reaction code with a full-kinetics confirmed this analytical results.

著者人数: 6名, 資料番号: AC1000020000

著者人数: 5名, 資料番号: AC1000019000

Starting loads observed in a free-jet type wind tunnel were investigated. Response functions between wall pressures and loads on a blunt model were derived by using Fast Fourier Transform (FFT). A large starting load was interpreted by the flow structure that separated nozzle flows, with Mach numbers less than the wind tunnel Mach numbers, impinge the model installed on a force measuring system (FMS). Probability functions for amplitudes of the starting loads were obtained to evaluate the possibility of excessive loads. The amplitude distribution of the loads and the experiencing frequencies were discussed from the viewpoint of normal random signals.

著者人数: 4名, 資料番号: AC1000028000

Prediction of harmful acoustic loading to payloads inside launcher fairing due to intense acoustic wave generated from propulsion systems at lift-off is an important design issue. Aero-vibro acoustics method developed in this study for predicting the acoustic loading consists of four elements. Hydrodynamics of jet flowfield generating aeroacoustic wave is computed by the high-fidelity Large-Eddy simulation. Computational aeroacoustics based on the full Euler equations is conducted to simulate propagating acoustic wave from the jet to the payload fairing. Then, finite element method is applied to simulate the structural vibration that radiates acoustic wave inside the fairing. Finally, acoustic behavior inside the payload fairing is also computed by the finite element method. An overview of the methods and recent work for validation and verification will be presented.

Acoustic test using Mach 2 cold jet was conducted for the H3 launch vehicle currently being developed in Japan. Effect of the clustered engines and the newly built movable launcher on the lift-off acoustics was investigated. The overall acoustic level taken by the farfield microphones did not show proportional increase in the number of engines, especially for the angles corresponding to the Mach wave radiation from the free jets. The observation here disagrees with the empirical prediction model. The computational fluid dynamics was also employed to analyze the acoustic mechanism of clustered engines.

An experimental study of scramjet nozzle was conducted to investigate how the nozzle flow is affected by ambient pressure. In order to elucidate the aerodynamic properties of nozzle flow, detailed measurements of thrust and wall pressure were carried out using cold nitrogen. Nozzle flow was also visualized using a shear sensitive liquid crystal. Wall pressure and shear stress distributions in an overexpanded nozzle showed that nozzle flow includes a crossing shock wave made at the side-fences. This flowfield can be approximated as a supersonic inlet flow compressed by side walls. The high-pressure region on the nozzle ramp generated by the shock waves results in a higher performance in scramjet nozzle than that estimated for a two-dimensiona l separation from the ramp. bient or "back" pressure in flight. The change in back pressure can significantly alter the thrust and pitching moment gen- erated in the nozzle, especially when the nozzle's internal flow is affected by external conditions. Thus, investigation of noz- zle performance in off-design conditions is important for es- timation of the feasibility of SERN for hypersonic vehicles. Both experimental and numerical studies on SERNs have been carried out to investigate the interaction between the nozzle internal flow and the external conditions. Simulated flight conditions in these studies have recently been extended to the hypersonic regime. Ruff in et al.4 calculated the flow- field in SERN in an underexpanded condition with an implicit Navier-Stokes solver and compared the results with experi- mental data. Watanabe5 investigated the flowfield in SERN in both underexpanded and overexpanded conditions exper- imentally. He also examined the effect of side-fence shapes on the flowfield. Zeutzius and Beylich6 conducted an exper- imental investigation on the effects of parameters such as aspect ratio and scarf angle on the flowfield in a scarfed SERN. These studies, however, concentrated on the investigation of the flowfield in SERNs, while the effect of the interaction, namely the separation, on SERN performance remains un- investigated. The objective of the present study is to investigate the variation of nozzle performance with the change in external conditions. A scarfed nozzle6 was selected to minimize the three dimensionality of the flow on the nozzle ramp. The effect of ambient pressure was measured in terms of nozzle pressure ratio (NPR), defined as the ratio of the nozzle total pressure to the ambient pressure. Total pressure of the nozzle flow was varied to change the NPR. Thrust delivered by a scramjet nozzle was measured in both underexpanded and overexpanded conditions. Obtained thrust efficiency was compared with predicted values. A simple prediction method based on the assumption that flowfield is two dimensional was adopted to predict the variation of nozzle performance. In the overexpanded condition, the treatment of interference of the internal flow with the external flow, viz., the separation of flow, becomes very important in the prediction of the noz- zle performance. In the present study, two-dimensional sep- aration was assumed. The flowfield in the nozzle in the over- expanded condition was observed to investigate the nature of the interference between the nozzle internal flow and the external conditions.

A three-dimensional calculation incorporated with the Baldwin-Lomax turbulent model was conducted to investigate internal flows in scramjet nozzles and the interference with external flows. The numerical results were compared with experimental results on the pressure, heat flux distributions and the friction loss. The numerical and experimental results revealed complex variations of nozzle performance by the crossing shock-waves and the secondary flow caused by the external flows. The calculations predicted the internal and the external loss within a discrepancy from the experimental results by about 4%.

Experimental research works for the ethanol propulsion system has been performed in JAXA/Space Transportation Mission Directorate (STMD) as a part of study for future reusable launch vehicle.. Under this program, 1000 kgf-class (2200 lbf) rocket engine testing, material-matching and critical heat flux (CHF) studies and other related works were conducted in 2009 to 2010. This paper provides an overview of this engine experimental program and the results.

A series of experiments was conducted for studying scramjet nozzles under Mach 8 flight conditions using a high temperature gas flow produced by combustion of monomethyl-hydrazine (MMH) and nitrogen tetroxide (NTO). The stagnation temperature was designed to be 3170K, and the stagnation pressure to be 1MPa. Scramjet nozzles with expansion ratios of 3 (EN3) and 5 (EN5) were examined. Delivered specific impulse (Isp) was found to be 17.3s (EN3) and 26.1s (EN5). Performances of nozzles were predicted using numerical codes, and energy release loss, kinetic loss and two-dimensional loss were identified. The two-dimensional loss was predicted to be 16.5% in the EN3 nozzle and to decrease to 7.0% in the EN5 nozzle. The kinetic loss due to chemical freezing was found to be about 4%. A friction loss evaluated from thrusts suggested laminar flow in the scramjet nozzle with the MMH/NTO experiments. Measured heat flux on nozzles also supported the laminar boundary layer. Experiments using a cold gas flow with the nozzles were also conducted to compare with the hot flow.

Combustion and vaporization of liquid fuels, under high pressure were investigated in the stagnation-point flow field. The diffusion flame was established slightly above the liquid heptane pool parallel to the surface, and the combustion mass flux was measured at various pressures up to 3.5 MPa. The mass flux increased with pressure, but it had a maximum around the critical pressure of heptane as was true with the burning rate constant of a droplet. In order to check the effect of chemical kinetics on this phenomenon, the vaporization mass flux was also obtained by using hot nitrogen instead of air for the stagnation-point flow. Since it showed the same tendency as in the case of combustion, it could be concluded that the maximum mass flux was caused by transport processes This is because properties such as thermal conductivity and diffusion coefficient, which control the transport processes, change significantly around the critical point, and therefore, chemical kinetics does not affect the present combustion phenomenon very, strongly. Flame extinction tests were also conducted in the same flow field, showing that the tretch rate (or the stagnation-point flow velocity) at extinction increased with pressure and attained a constant value but decreased again around 2.0 MPa. This tendency did not depend on the kind of alkane fuels and the critical pressures. Extinction data for alkanes were very similar to those for methane as predicted for heptane by Seshadri.

43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Cincinnati, OH, July 8-11, 2007, 資料番号: AC1000093000, レポート番号: AIAA-2007-5389

This report presents recent research activities of the Combined Propulsion Research Group of Japan Aerospace Exploration Agency. Aerodynamics and combustion of the scramjet engine and the rocket–ramjet combined-cycle engine, structure and material for the two engines and thermo-aerodynamic of a re-entry vehicle are major subjects. In Mach 6 condition tests, a scramjet engine model produced about 2000 N net thrust, whereas a model produced thrust almost equal to its drag in Mach 12 condition. A flight test of a combustor model was conducted with Hyshot-IV. A rocket–ramjet combined-cycle engine model is under construction with validation of the rocket engine component. Studies of combustor models and aerodynamic component models were conducted for demonstration of the engine operation and improvement of its performances. Light-weight cooling panel by electrochemical etching examined and C / C composite structure was tested. Thermo-aerodynamics of re-entry vehicle was investigated and oxygen molecular density was measured also in high enthalpy flow.

A quasi-one-dimensional analysis with either chemical equilibrium or finite-rate reaction was conducted to evaluate the effects of contamination due to the vitiation on engine performances, namely thrust production. The incoming flow state was calculated based on measured values, 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 2/3 of the reduction in the thrust production with vitiation observed in the engine tests. With finite-rate reaction calculation, possibility of a further reduction due to contamination effect on combustion mechanism was found, which was responsible for the 1/3 of the measured reduction. The one-dimensional analysis was further pursued to find a 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.

Quasi‐one‐dimensional analyses with either chemical equilibrium or finite‐rate reaction were conducted to evaluate the effects of contamination from vitiation on engine performances, namely thrust production. The incoming flow state was calculated based on measured values, and a higher total enthalpy for the freestream through a vitiation heater compared to that through a storage heater was found, which reduced the thrust production by about 10% in equilibrium case. With finite‐rate reaction calculation, possibility of further reduction of the thrust production by about 5% due to contamination effect on combustion mechanism, was found.

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit(11-14 July 2004), Fort Lauderdale, Florida, 資料番号: ARDS04109000, レポート番号: AIAA 2004-3341

39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, Alabama, July 20-23, 2003, 資料番号: AC1000095000, レポート番号: AIAA-2003-4739

38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Indianapolis, Indiana, July 7-10, 2002, 資料番号: AC1000097000, レポート番号: AIAA-2002-4129

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 37th, Salt Lake City, UT, July 8-11, 2001, 資料番号: AC1000098000, レポート番号: AIAA-2001-3202

Aerodynamic performance of scramjet engines was measured by using 1/5-subscaled (0.42 m-long) models of H 2-fueled scramjet engines. The air capture ratio, the drag and the wall pressure distributions in the engines were measured using the models. By considering the momentum balance between the inlet and the nozzle, the internal drag of engines were expressed with the air capture ratio, pressure recovery factor and the total temperature ratio. To evaluate the pressure recovery based on the engine drag, the total pressure loss was analytically estimated as the sum of pressure losses for oblique shock and friction in the inlets. The friction loss in engine internal flow was evaluated by one-dimension al analysis of duct flow with changes in the cross-sectional area and total temperature. It was found that the pressure recovery across the engine was about 6Q% in the Mach 4 engine without struts and 19% in the Mach 6 engine with a small strut. With a larger strut, the pressure recovery was reduced to about 10% in the Mach 6 condition. This evaluation of the pressure recovery enabled calculation of the flow condition at the entrance of the combustor for evaluation of the engine performance in mixing-controlled or reactioncontrolled combustion. Nomenclature