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This paper discusses a flush air-data sensing (FADS) system for a sharp-nosed hypersonic vehicle designed to estimate flight air data in real time at hypersonic speeds. The design’s target condition is Mach 5.0 to 7.0 with an angle of attack within 5 deg. The FADS system estimates air data by relating them to surface pressures measured from surface-mounted ports on the vehicle forebody. Unique combinations of pressure ratios were used to estimate the freestream dynamic pressure, which was a primary parameter for the flight experiment, the angle of attack, and the Mach number. The proposed FADS estimation algorithm was validated through numerical simulation, which was also used to generate datasets of surface pressures for given flight conditions. To handle possible sensor errors related to estimation accuracy in real-time estimation for a flight experiment, redundant systems were implemented. The results indicated that the designed FADS system can estimate air data within the uncertainty of 4.8% for a single estimator by considering sensor errors for freestream dynamic pressures in the range of 0–100 kPa, including targeted and off-design flight conditions. The proposed algorithm can estimate the air data with an acceptable level of uncertainty while retaining the robustness of estimation to sensor failures with low computational cost.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-08-31, 資料番号: PA2110005000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-07-09, 資料番号: PA2110004000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2018-01-19, 資料番号: PA1810076000

形態: 図版あり, Physical characteristics: Original contains illustrations, Accepted: 2016-06-11, 資料番号: PA1710041000

To reduce the cost of space transportation, air-breathing engines are considered to be candidates for propulsion. However, to cover a wide range of flight speeds, the propulsion system has to operate in various modes to be efficient under incoming atmospheric-air conditions. The Japan Aerospace Exploration Agency is proposing a rocket-based combined cycle engine for operation under various condition, an ejector-jet mode being adopted for the low-speed regime. The suction performance ejector-jets has long been studied experimentally and numerically at JAXA, and little success has been achieved in explaining the deterioration of suction performance with high-temperature gas or light gas such as helium. In the present study, based on former models, a simple one-dimensional model was introduced incorporating the mixing effects of the primary flow (rocket flow) and secondary flow (induced air flow). The results were compared using several experimental and numerical data to check the plausibility of the model. It was found that if greater mixing occurs, suction performance is degraded, explaining the actual phenomena of the experiments., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1610077000

資料番号: ARDS06125000

著者人数: 3名, 資料番号: AC1000036000

condition was calculated by incorporating results of a chemical equilibrium numerical code. The fuel rates causing the flow choking were successfully reproduced and matched with the limit of fuel rates observed in our tests. Inlet unstart of the engine caused the loss of the ejector-pumping effect in the diffuser system, which triggered flow choking. Choking was also promoted by the drag of the gas sampling rakes. The choking in diffuser flow and the engine unstart may be coupled, resulting in hysteresis in the diffuser breakdown, which was also experienced in our tests. As the rocket-based, combined-cycle engine is scheduled to be tested under the Mach 4 condition in 2007, the safety of the experiment is now a primary issue. The current analysis showed that the new engine would easily cause choking of diffuser because of the larger propellant supply rates and the relatively low specific impulse. Reasonable operation of the wind tunnel to control the flow choking in future tests was examined.

Combination of a air-breathing engines and rocket engines (termed as Rocket Based Combined Cycle engine) gives an opportunity to reduce onboard oxygen consumption and to increase system weight margins for various application from cruiser to launch vehicles. For launch vehicle applications, combination of a ramjet/scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines was proposed (termed as Rocket-Ramjet Combined Cycle engine, T. Kanda, et. al., JPP, 19(2003), 859), and related R&D activities are undergoing at Japan Aerospace Exploration Agency (JAXA), Kakuda Space Center targeting hydrogen as the fuel due to its high Isp and cooling performances. Use of hydrogen fuel, on the other hand, can be costly due to the fuel cost and difficult-to-handle nature of liquefied hydrogen. Thus, hydrocarbon fuel such as ethanol was under consideration for the application to the future reusable launch vehicles termed as the ‘reference system,’ targeting manned operation to LEO. This presentation is to show the status of system analysis, flight demonstration plan, and related research efforts.

Ramp-compression-type inlets for the combined cycle engine were investigated in a Mach 4 wind tunnel. Geometries of the inlet models were changed to investigate their effects on the aerodynamic performances and the starting characteristics. A bent cowl improved the starting characteristic by weakening the shock from the cowl leading edge to suppress the shock-induced separation and by reducing the internal contraction ratio of the cowl duct, thus, avoiding choking at the inlet throat. With the current geometries, the inlet started with a slightly larger internal contraction ratio than that restricted by the Kantrowitz-Donaldson limit. The ratio of the height of the cowl duct of the inlet to the incoming boundary layer was found to affect the starting characteristic. The model with a higher capture ratio showed better total pressure recovery performance.

著者人数: 3名, 資料番号: AC1000035000

2005 Asian Joint Conference on Propulsion and Power (January 27-29, 2005.), Kitakyushu, Fukuoka, Japan, 資料番号: AC1000138000, レポート番号: AJCPP2005-22132

著者人数: 5名, 資料番号: AC1000026000

著者人数: 5名, 資料番号: AC1000025000

Combustion tests of a scramjet engine were conducted to obtain thrust, lift, and pitching moment essential for evaluation of the engine’s basic performance and for determination of the requirements to control the system of integrated engine and body. The engine was a side-compression type with a swept-back angle of 45 deg, and the inlet was open on the cowl side. Thrust, lift, and pitching moment were measured safely by a force measuring system under Mach 4, 6, and 8 simulated  ight conditions including engine unstart. Measurements changed with transition of combustion conditions.Thrust decreased and lift increased suddenlywhen the engine fell into unstart conditions in theM6 andM8 tests although the changewas gradual in theM4 tests. Effects of engine conŽ gurations were also investigated. Installation of a full-height strut improved combustion performance but increased drag more than thrust.

著者人数: 4名, 資料番号: AC1000023000

著者人数: 5名, 資料番号: AC1000022000

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.

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

Joint Propulsion Conferences, 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference（July 14-17, 2013.), San Jose, California, United States., Ground firing-tests of rocket-ramjet combined-cycle engine model with 3 m length was conducted under Mach 8 flight condition at Ramjet Engine Test Facility in Kakuda Space Center of JAXA. Numerical analysis was also conducted for the engine under Mach 8 flight condition. The effects of rocket chamber pressure and a secondary fuel injection on the engine performance such as thrust and specific impulse were examined., 資料番号: AC1400049000, レポート番号: AIAA2013-3668

Engine system analysis upon a Rocket-Based Combined Cycle (RBCC) engine for the booster stage of a TSTO launch vehicle was conducted, to evaluate the effects of hydrocarbon fuel type upon the performance of the engine and the whole launch vehicle. A higher thrust and cooling performance with methane was shown, while some technical challenges were also noted. Combination of the scramjet flow pass and the rocket engines, i.e., rocket engine being embedded into the scramjet flow pass or separately mounted to the airframe, was varied to show that the separated configuration mitigated cooling problem, and resulted in a better performance for mid-speed regime., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1610081000

In this study, numerical analysis was conducted for the scramjet mode operation of a JAXA rocket/ramjet combined cycle engine called E3 under Mach 8 flight conditions. The objective of the study is to examine the effects of rocket chamber pressure Pc and secondary fuel injection on the engine performance and related flow physics before the engine combustion tests. The calculation results revealed that the engine produced higher specific impulse Isp as well as thrust with higher Pc. However, the net thrust acting on the internal ramjet duct wall was smaller than the embedded rocket thrust. This is due to the fact that the rocket exhaust increased the friction drag on the combustor walls over the thrust due to combustion with the captured air. On the other hand, the secondary fuel injection in the diverging combustor had almost no impact on the thrust augmentation. This partially comes from that at lower Pc the combustion efficiency was much lower than expected because of the lower temperature and larger amount of H radical production in the combustor air flows. In addition, an auxiliary fuel injector was tested to examine the effects on combustion enhancement for the rocket residual fuel. A small thrust gain was found by using this injector.

A rocket-ramjet combined cycle engine model, embedding twin rocket chamber on top wall side of a scramjet flow-pass, was tested under ramjet- (dual-) mode operation at M6 flight conditions. The rocket chamber was driven with gaseous hydrogen and oxygen at nominal operation condition of 0.6 MPa in chamber pressure and 6.0 in mixture ratio. Gaseous hydrogen was also injected through injector orifices within the ‘ramjet’ combustor to pressurize the combustor. A simple, one-dimensional calculation was conducted based on wall pressure measurement to evaluate combustion mode and engine performance. Engine exhaust samplings as well as Pitot pressure measurements were conducted to evaluate the calculation results. Different combustion modes were observed, which were found to have sizable effects on the engine performance.

A free-piston shock tunnel (FPST) is one of the most useful ground testing facilities for hypervelocity flow research of re-entry vehicles and scramjet engines. For an efficient operation with tuned piston motion, the design of facility and the comprehension of the physical phenomena in a FPST, a numerical simulation which can properly predicts the flow with actual losses is required. But there are few successful numerical methods which can simulate its overall performance. In the present study, numerical method was developed by using the KRC shock capturing scheme and by modeling the flow losses in suitable forms for a quasi-1D numerical computation. The present numerical results were compared with the data obtained in two different facilities, T4 and T5. The applicability of the present numerical method as a design tool is discussed briefly.

An experiment was conducted to investigate the effect of injector/combustor geometry on combustion-induced peak wall pressure and associated upstream influence, as well as on mixing/combustion characteristics at an entrance Mach number of 2.5. The length of the constant area section downstream of injection orifices had a strong influence on the above-mentioned characteristics. However, the sweep of the rearward-facing steps on both side walls had little effect on these characteristics, nor did reversing them have any effect. The peak wall pressure and the length of the upstream influence agreed qualitatively with predictions of an analytical model and an empirical formula developed at Johns Hopkins University. Fuel jets injected from the model with the longest constant area section and the fuel equivalence ratio of unity, coalesced at a very early stage downstream of the fuel injection orifices. This coalescence led to a decrease in mixing rate downstream, despite the higher degree of mixing near the injection orifices. The combustion efficiencies were higher than those obtained at NASA Langley in the upstream region due to the higher mixing rate near the injection orifices.

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.

In order to get experimental data in subsonic Mach number flight region, two small hybrid rockets installed the ejector-jet were launched, and then the flight analysis was performed. Numerical simulations were conducted for clarification of the flowfields and prediction of the ejector-jet performance. The CFD results revealed the compound compressible flow phenomena involving oblique shock waves and the subsonic flows. The effects of the flight Mach numbers and the rocket pressures to the suction performances were also investigated.

A combustion-capable combined cycle engine model which was constructed based on the rocket and ramjet technology was tested in Mach 4 flight condition. At this speed, engine is designed to shift its operation mode from an ejector-jet to a ramjet. Both modes were simulated by changing the rocket combustion pressure. Even with full rocket exhaust, no effect to the air flow could be observed. The injection point of the secondary fuel affected thrust performance. In the ramjet mode, the pressure rise due to the fuel combustion traveled to the entrance of the combustor, but it stayed near the injection point in the ejector-jet mode.

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.

44th AIAA Aerospace Sciences Meeting and Exhibit(9 - 12 January 2006), Reno, Nevada, 資料番号: ARDS05397000, レポート番号: AIAA 2006-1027

44th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 9-12, 2006, 資料番号: AC1000085000, レポート番号: AIAA-2006-0224

44th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 9-12, 2006, 資料番号: AC1000084000, レポート番号: AIAA-2006-0223

43rd AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 10-13, 2005, 資料番号: AC1000080000, レポート番号: AIAA-2005-1045

43rd AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 10-13, 2005, 資料番号: AC1000082000, レポート番号: AIAA-2005-1210

著者人数: 6名, 資料番号: AC1000017000

39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, Alabama, July 20-23, 2003, 資料番号: AC1000096000, レポート番号: AIAA-2003-4737