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Skin-friction values on the surfaces of clustered nozzles were measured using a pointwise local skin-friction sensor. This sensor has a floating head that can detect the local shear force using the flow passing over the head surface. Cold-flow wind-tunnel experiments were carried out with clustered nozzle models. The test models had three or four clustered cell nozzles, or a nonclustered two-dimensional nozzle with a designed Mach number of 3.5, followed by a straight planar section. The experimental conditions were chosen according to the cell nozzle jet expansions. Hence, some conditions involved an oblique shock wave impinging on the bottom surface and others did not. The measured local skin-friction coefficients were compared to those predicted by using an existing prediction model. The results showed that the local skin-friction value significantly increased when the cell nozzle expansion was under the overexpansion condition, whereas the measured values were identical to those predicted by the existing model under the underexpansion condition. The physics behind this observation showed that the expansion waves hardly affected the skin friction, but the oblique shock wave and crossing neighboring jet boundaries could increase the skin friction., 資料番号: PA1510020000

Spike surface flowfields and base wake behaviors of clustered linear aerospike nozzles under the presence of external flow were characterized using surface pressure measurements, flow visualizations using the background-oriented schlieren method, and theoretical modeling. The test model consisted of three clustered cell nozzle modules or a nonclustered two-dimensional nozzle with an exit Mach number of 3.5, followed by a straight section and a contoured spike. The model was exposed to an external flow of Mach 2.0 to simulate off-design transonic flight conditions. The measured surface pressure distributions indicated that the entire flowfield on the spike surface became more two-dimensional under the presence of external flow. Periodic compressions and expansions of the cell nozzle jet were eliminated, but the pressure continued to increase downstream. The results of the flow visualizations found that this was mainly governed by the external flow that controls the jet expansion by forming a slip surface. This mechanism also made the spike base pressure higher than environmental pressure. The physics-based model predicted the measured surface pressure distribution well and explained the mechanism behind the influence of external flow. The effects of the external flow showed improvements in thrust performance in off-design flight conditions., 資料番号: PA1510019000

The influence of external flow on the plume physics of cell nozzle flows and the cell base wake behavior of simplified clustered linear aerospike nozzles were characterized using surface pressure measurement, flow visualization using the background oriented schlieren method, and theoretical modeling. The test model comprised three or four clustered cell nozzles, or a nonclustered two-dimensional nozzle with a designed Mach number of 3.5, and was connected to a straight section followed by a straight ramp inclined at 12 deg. In the cases with or without external flow, the measured pressure distributions on the surface of the test model were found to be different in the presence of external flow: The ramp surface pressures were maintained at a certain level above the environmental pressure according to the cell nozzle flow conditions, whereas the pressure distributions in the cases without external flow demonstrated highly periodical compression/expansion features. The results of flow visualization indicated that the pressure maintained on the ramp surface depends mainly on how the external flow controls the jet expansion by forming a slip surface. The cell base wake was found to behave as a closed wake in the presence of external flow. Those pressure behaviors were successfully predicted using the physics-based theoretical models established in this study., 資料番号: PA1510018000

For space plane with air-breathing engines, equipping Single Expansion Ramp Nozzle on its aft-body is beneficial for thrust augmentation. The cowl of this nozzle is truncated (thus, termed as an external nozzle), in order to reduce its weight, friction drag, and cooling requirement, and also to attain pressure recovery (so-called altitude compensation) on the ramp wall by the impingement of pressure waves generated by the interference between external nozzle flow and ambient flow on ramp wall surface, in case of over-expanded condition. The present study is to evaluate the effects of ambient flow on the external nozzle flow field, as little research has been conducted to elucidate the effects. Room temperature nitrogen gas was injected to simulate air-breathing engine exhaust in cases without and with ambient flow at a supersonic semi-free jet wind tunnel. A straight expansion ramp was employed as the external nozzle. Pressure distributions on the ramp wall were measured and results without and with ambient flow were compared. Comparison showed that presence of ambient flow reduced the magnitude of ramp wall pressure induced by the impingement of pressure waves, and thus, the strength of pressure waves generated by the interference. Flow characterization by pressure wave tracking showed that spatial variation in ambient flow pressure caused this weakening of the incident pressure wave upon the ramp surface. Thus, altitude compensation effect in the over-expanded condition became smaller than that without ambient flow., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1610046000

The influences of the molecular weight and total temperature of a primary gas on the flow fields of constant-area mixing tubes for ejector-jets were investigated. First, in order to investigate the exit choke, parametric studies were carried out for the exit-choke phenomena in the constant-area mixing tubes by using a one-dimensional model. The effects of the incoming Mach number, wall friction, expansion ratio of a primary nozzle, and specific heat ratio on the exit-choke phenomena were reported. Moreover, flow fields having primary flows with high total temperatures were investigated by CFD. The transition points from the exit-choke mode to the free-jet choke mode obtained by CFD were in good agreement with those obtained by the one-dimensional model. The pressure recovery performances in both choke modes were compared. The decrease in the static pressure at the exit of the subsonic diffuser attached downstream of the mixing tube was observed.

A new variable termed the ejector drive parameter was proposed as a dominant parameter of the pressure recovery ratio of constant-area mixing tubes. It was derived from the one-dimensional analysis of a zero-secondary flow mode. This parameter represented the strength of the normal shock wave generated in the rocket exhaust that is spreading to the entire cross section of mixing tubes under the zero-secondary-flow condition. Parametric studies were carried out for investigating the pressure recovery ratios of the constant-area mixing tubes, and the analytical results were plotted against the ejector drive parameter or several other parameters. It was determined that the influences of each parameter on the pressure recovery ratios were minimized by using the ejector drive parameter. In order to validate the results of the parametric studies, cold flow experiments were conducted by using supersonic-subsonic ejectors with cylindrical or rectangular constant-area mixing tubes. The measured pressure recovery ratios were in good agreement with the pressures predicted by the one-dimensional analysis.

Flow fields in the constant-area mixing tubes of ejector jets were investigated under the starting-limit conditions of an aerodynamic choking mode by performing numerical simulations and cold flow experiments. Pressure recovery was almost completed in the shock-train region. The length of the shock-train region (Lst) was measured under various conditions. Lst was proportional to the mass flow rate ratio of the secondary flow to the primary flow when this ratio was less than 0.15. On the other hand, Lst became almost constant when the mass flow rate ratio exceeded 0.15. Numerical studies showed that this change was caused by the difference in the mechanism of the flow fields. In the cases with low air mass flow rates, the primary and secondary flows almost mixed in a region between the inlets of the mixing tubes and the choking points. The pressure was recovered by a pseudo-shock-wave generated downstream of the choking point. On the other hand, when the mass flow rate ratio was higher than 0.15, the primary and secondary flows were clearly separated at the choking point. The pressure recovery was achieved by the mixing between the primary and secondary flows downstream of the choking point.

資料番号: ARDS06125000

Breakdown of diffuser flow was often observed in our scramjet engine tests. This facility operation may damage the engine wind-tunnel and should be prevented. An one-dimensional analysis was applied to the diffuser flow to identify the causes of the flow breakdown. All the losses and gains by engine and friction loss in the diffuser were represented by point-sources of mass, momentum and energy. The thermal choking condition was calculated by uses of a chemical equilibrium code. The fuel rates causing the flow-choking successfully reproduced the limit fuel rates observed in our tests. Inlet-unstart of engine lost the ejector-pumping effect in the diffuser system to trigger the flow choking. The choking was also promoted by the drag of the gas sampling rakes. The choking in diffuser flow and the engine unstart may couple each other to cause hysteresis in the diffuser breakdown, which was also experienced in our tests. A rocket-based, combined-cycle (RBCC) engine will be tested under the Mach 4 condition. The engine easily causes the choking of diffuser because of the large propellant supply rates and the relatively-low specific impulse. Operation of the wind-tunnel was discussed to control the flow choking in the tests.

It is necessary for realizing the space-plane with air-breathing engines to develop the performance prediction tool for each component of this engine, which estimates the performance with the sufficient accuracy and much lighter calculation cost than 3D-CFD. The objective of the present study was to predict the thrust performances of a clustered external nozzle by experiment based prediction model (not CFD). Two-dimensional wave model was developed through our previous experiments. Experimental validation of this model showed that three-dimensionality of the nozzle flow, resulting from clustering, became more remarkable, hence, the prediction errors of nozzle thrust performances also became larger with increase of flight altitude. The effects of clustering on thrust performances were experimentally investigated through the element tests. Finally, the clustering effects could be included into the above-mentioned model by modifying the model input value, e.g., averaging the flowpath exit pressure (after the rapid expansion) and cell base pressure. In conclusion, this model predicted the nozzle thrust performances within an error of 10 % and calculation time was around 1 minute for each condition.

A simple variable Mach number nozzle for aerodynamic engine tests was designed and calibrated. In this new nozzle, only one hydraulic actuator made freestream Mach number change from 2 to 4 during a run of the wind tunnel. The optimization of the nozzle contours was conducted by the 2-D computational simulation. The CFD results predicted that the deviation of Mach number from the averaged value was less than 1% and flow angle was less than 1 degree in the test region. The results of the calibration tests of the nozzle showed that the deviations of Mach number distributions in the test core (60mm×60mm) were less than 1.5% from Mach 2 to 3.5 and 3.5% at Mach 4. This result assured that this nozzle was suitable for aerodynamic engine tests. The aerodynamic tests of scramjet engine models in acceleration/deceleration conditions were carried out with this nozzle. Air capture ratio and internal drag were measured. When freestream Mach number decreased, the soft unstart in the inlet was observed around Mach 3. It resulted in increase of the internal drags.

A method to evaluate aerodynamic performances of scramjet engines by using multi-probe rakes was proposed. The aerodynamic tests were carried out under Mach 4 flight conditions. The Pitot and static pressures were measured at 250 points in the cross sectional area of the engine exit by the rakes. Local mass flux and thrust function were evaluated from the pressure measurement at each point and integrations of these values enabled to obtain the mass flow rate and the stream thrust at the engine exit. The air capture ratios were independently measured by the rakes and a conventional choked flowmeter. The air capture ratios measured by these two methods agreed within 2%. It was found that the rakes enabled to measure the air capture ratio more simply than the flowmeter. Additionally, the effect of boundary layer ingestion to an internal drag was investigated by the rakes. The decrease of air capture ratio measured by the rakes showed that the ingested boundary layers were separated in the inlet. The pressure drag of inlet increased by the separation and the pressure thrust decreased by the decrease of air capture ratio. As a result, the internal drag increased when the forebody boundary layer was ingested.

Experiments and a quasi-one-dimensional equilibrium calculation were conduced to estimate performances of a dual-mode combustor with ethylene fuel injection in a directconnect wind-tunnel facility in JAXA Kakuda Space Center (JAXA-KSPC). In the facility, a gaseous H2 fueled lean combustion heater was used to make high enthalpy flow. The high enthalpy flow was then accelerated with a Mach 2.5 supersonic nozzle. The experiments were executed with a rectangular dual-mode combustor with constant area section and following diverging section at various fuel equivalence ratios. Fuel injectors upstream of cavity flameholders were installed in the constant area section and the diverging section on opposite walls. Heat flux distributions in the streamwise were measured using heat flux transducers mounted flush with the combustor inner wall. The compositions of gas at the exit cross-section of the combustor were sampled using a water-cooled reaction-quench sampling probe. The compositions of sample gas were eccentrically located on the fuel injection side under both of scramjet and dual-mode operation. Therefore, there is a need to consideration of the non-uniformity to estimate state quantities like combustion efficiency and heat flux by quasi-one-dimensional calculation. Heat flux distributions calculated using the “two-flow tubes model” indicated good agreement to measured heat flux. The combustion efficiency defined as the fraction of the oxygen rate that reacted over the total oxygen flow rate was closer to the experimental value than that defined as the fraction of the reacted fuel rate.

Base wake behaviors of an aerospike nozzle under the presence of external flow were characterized by surface pressure measurements, flow visualizations using the background oriented Schlieren method, and theoretical modeling. The test model consisted of three clustered cell nozzle modules or a nonclustered two-dimensional nozzle with a designed Mach number of 3.5, followed by a straight section and an ideal contoured spike with truncated lengths. Additionally base blocks which had a 60, 75, and 90 inclined trailing edge against the flow direction were used to investigate the effects of base angles and depths on the spike base pressure distributions. The test model was exposed to an external flow of Mach 2.0 to simulate off-design transonic flight conditions. Results of flow visualizations and pressure measurements showed that the base pressure could be higher than the environmental pressure level under the presence of external flow with the base angle of 60 model. The tendency was stronger when the base depth was shallower. The physics-based model to predict the base pressure predicted the measured base pressure distribution well.

AIAA SciTech 2014, 52nd Aerospace Sciences Meeting （January 13-17, 2014. National Harbor), Maryland, United States., Spike surface flowfield and base wake behaviors of clustered linear aerospike nozzles under presence of external flow were characterized by the wind tunnel experiment with surface pressure measurements and a flow visualization using background oriented schlieren technique. The test model has three clustered cell nozzle modules or nonclustered two-dimensional nozzle with a designed Mach number of 3.5, and is connected to a straight section following a contoured spike. The model is given with a Mach 2.0 air stream as an external flow to simulate an off-design transonic flight condition. Measured pressure distributions on the spike surface indicated that the external flow makes the entire flowfield on the spike surface more uniform compared to those for the conditions without-external flow by eliminating periodic compression/expansion feature of cell nozzle jet. Results of flow visualization found that those tendencies are mainly governed by the fact that the external flow controls jet expansion by pressing it against the wall by forming a slip surface between cell nozzle flow and external flow. This fact made the spike base pressure higher than environmental pressure level when the external flow exists. Those effects of external flow showed advantages in thrust performance operated at off-design flight conditions., 資料番号: AC1400046000, レポート番号: AIAA2014-0551

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.

Large-amplitude load oscillations were observed during the tests of a hypersonic engine model in a freejet-type wind tunnel. To clarify the cause of the oscillations and their characteristics, oscillating wall pressures and loads on a drag model and engine models were investigated. Flowe eld was observed by shadowgraph to determine the cause of the large starting loads. Power spectral density functions and probability functions of wall pressures and loads were derived by the fast Fourier transform. The amplitude of the unsteady frontal pressure was correlated with the dynamic pressure. The magnitude of the starting load was related to the drag coefe cient of the models, and the expected maximum peak loads of a large-scale ramjet engine test facility were evaluated. Engine unstart loads were also simulated by means of secondary e ow injection into a small-scale model of a ramjet engine. With these methods, characteristics of engine unstart loads and the possibility of sensing engine unstart in its early phase were studied. Engine unstart could be sensed with pressure measurement around the engine throat before it became severe. Furthermore, engine unstart loads associated with scramjet engine combustion were related to the drag coefe cient of the engine.

Flow interaction between external flow and cell jet flows of a simplified clustered linear aerospike model was experimentally investigated based on the surface pressure measurements and a flow visualization using background oriented Schlieren (BOS) technique. The test model has four clustered cell nozzles with a designed Mach number of 3.47 and is connected to a straight section following a 12-degree-inclined straight ramp. It is set up in a Mach 2.0 supersonic wind tunnel to simulate conditions of supersonic external flow impinging on the cell nozzle jets and ramp surface. Under conditions with or without external flow, pressure distributions on the ramp surface were measured and compared to flow visualization results. Results of measured pressure distributions on the test model surface for the with-external flow case were obviously different due to the external flow: ramp surface pressures were maintained at a certain level above the environmental pressure level according to the cell jet expansion conditions, and the pressure distribution for conditions without-external flow can be nearly approximated by the method of characteristics. Furthermore, the external flow makes the entire flowfield on the ramp more uniform compared to that for the conditions without-external flow by eliminating periodic compression/expansion feature of jet. Flow visualization found that those tendencies are mainly governed by the fact that the external flow controls jet expansion by pressing it against the ramp wall surface. While flowfield was investigated, BOS technique itself was assessed to have better sensitivity and accuracy for a given setup.

A Relation between flow structure and local skin friction on the wall surface of the clustered linear aerospike nozzle models was investigated by conducting cold flow experiments using a skin friction sensor which has been developed in this study. The skin friction sensor is capable of measuring the local skin friction value on the wall. Measured skin friction value was compared to some existing prediction model and results found that the local skin friction value becomes drastically different from the one predicted by existing models. In cases of over-expansion condition skin friction was drastically increased by oblique shock wave. In cases of under-expansion condition skin friction was increased which was caused by interference of jet boundary of cell-nozzle flows jetted from adjoined cellnozzle.

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.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1710076000

Effects of design parameters on fuel-jet / airflow mixing in presence of pseudo-shock system were experimentally investigated. The fuel-jet was simulated by inert helium injection, and the pseudo-shock wave system was generated in a duct by backpressurizing Mach 2.5 airflow with a mechanical valve. Cross-sectional aspect ratio of the duct, injection conditions such as orifice diameter and injection pressure, and pseudo-shock wave location with respect to the injector location were varied. Crosssectional aspect ratio had sizable effects on mixing characteristics through interaction of injectant plume with wall and/or boundary (i.e., neighboring jet). Injectant mass flow rate was dominant among other injection parameters, and a length scale taking mass flow rate into account was found to be effective to summarize mixing data. Location of the pseudo-shock wave system had sizable effects on mixing characteristics even the shock train itself was terminated at the injector location, and better mixing was attained for closer location.

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.

*, † ‡ # An experimental investigation was conducted to reveal jet penetration and mixing performance of pulsed injection in a Mach 2.5 crossflow. Helium and nitrogen gas were injected perpendicularly through flush-mounted circular sonic orifice. Probing techniques including species composition sampling and high speed framing schlieren were employed to determine the penetration and mixing performance at several downstream locations. Our investigation consisted essentially of two parts. The first part was an investigation of the continuous jet. The performance of the continuous jet was mainly controlled by effective velocity ratio (r) as the square root of the momentum-flux ratio and the orifice diameter (d). The centerline trajectories of the jets and the maximum concentration decay were collapsed by the rd scale and the ratio of oncoming boundary-layer thickness to the injector diameter. The second part was an comparison of the performance between the pulse and continuous jets. The penetration of the pulse jet was adjustable by changing the pulse duty cycle at a condition of fixed injectant mass flow rate. Even at a condition of fixed injection pressure, the pulsed injection showed better mixing performance and the higher penetration, due to the fluctuation of the large-scale eddies in the jet associated with the fluctuation of the bow shock in front of the jet.

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

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

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

Aerospace Sciences Meeting and Exhibit, 38th, Reno, NV, Jan. 10-13, 2000, 資料番号: AC1000076000, レポート番号: AIAA-2000-0616

A model scramjet engine was tested at the NAJLKRC RamJet engine Test Facility (RJTF). The engine had an ignition system consisting of plasma torches, pilot fuel injectors, and backward-facing steps upstream of the main fuel injectors. Tests were conducted under conditions of flight Mach numbers of 4, 6 and 8 at altitudes of 15, 25 and 35 km respectively. Tests results showed that the ignition system had sufficient ability to attain intensive combustion in the engine in the Mach 4 condition, whereas it lacked the ability in the Mach 6 and 8 conditions. The ignition system was thus modified and then tested at the Mach 6 condition. At first, a recess with auxiliary injectors was placed downstream of the step to obtain shorter ignition delay and longer residence time than those in the step-base region. Oxygen was injected through the auxiliary injectors to imprtve the fuel-rich condition within the recess. As a result, intensive pilot flame was attained, but ignition of the main fuel occurred only under limited conditions. To attain higher pressure level around the ignition system, the recess was removed and a short strut was mounted on the top-wall. Consequenty, intensive combustion of the main fuel was achieved over a wide range of operational conditions. Using an ignition parameter, defined as pressure-length products divided by velocity (ph/u), the ignition limit was expressed on a temperature-ignition parameter diagram. Introduction The supersonic combustion ramjet (scramjet) is expected to be the most effective propulsion system for the Single Stage To Orbit (SSTO) transportation vehicles and hypersonic transportation vehicles of the next generation. Many studies on scrramjet components such as inlet, combustor and nozzle, have been carried out to obtain a better understanding of the performances of the individual components. *~4 However, it is expected that intensive interactions among these components will occur in real engines. For example, a rise in pressure due to combustion causes separation which propagates upstream * Researcher, Ramjet Combustion Section, Ramjet Propulsion Division. Member ALAA. t Researcher, Ramjet Control Systems Section, Ramjet Propulsion Division. $ Researcher, Ramjet Performance Section, Ramjet Propulsion Division. § Project Engineer, Engineering Section, Engine & Control Equipment Department. Copyright © 1996 by the American Institute of Aeronautics and Astronautics. Inc. All right reserved. into the inlet and changes the.inlet back pressure, which may result in an unstart condition in the inlet. In the case of such interactions, the total performance of the engine cannot be evaluated based on a linear combination of the obtained performances of the individual components. Thus, tests of the models of whole engine are necessary to elucidate the interactions among these components and the overall performance of the whole engine. Testing of the models of whole engine requires rather big and expensive wind-tunnel facilities. Thus, only a limited number of results on whole engine tests have been reported.^ Using the test facility (denoted as RamJet engine Test Facility, RJTF) at NAL-KRC, we have conducted tests of a model scramjet engine in various flight conditions. In flight conditions of Mach 4, intensive combustion was attained with an installed ignition system.^ However, in flight conditions of Mach 6 and 8, it was found that intensive combustion were not attained in the model with the original ignition system.8>9 To improve ignition ability, supplemental equipment such as a recess with auxiliary injectorsS) or a short strut *-" were installed in the model engine. In the present paper, ignition characteristics of the engine with and without these modifications are reported. Test Facility and Engine Model Test facility The RJTF was designed to ssimulate flight conditions of M4, M6, and M8. The facility is equipped with two heating systems to obtain high enthalpy flow. A storage air heater (denoted as SAB) is used to obtain pure hot air for the M4, M6, and M8 flight condition tests. Two types of vitiation air heaters (VAH) fueled by H2 are used to obtain vitiated hot air for the M6 and M8 flight condition tests, respectively. The M6 flight conditions are achieved with the SAH alone or the VAH alone (denoted as M6S and M6V conditions, respectively), whereas the M8 conditions are achieved with a combination of the SAH and the VAH. Table 1 shows the operational conditions of the RJTF and the simulated flight conditions. The high enthalpy flow obtained at the heaters was accelerated through rectangular nozzles. The exit cross section of the every nozzle was 510 mm x 510 mm. The engine model was set in a low pressure chamber which was connected to a steam ejector system to reduce the pressure within the chamber for simulation of high altitude atmosphere. The engine model was located in such way that it breathed in the boundary layer in the facility nozzle. The displacement thickness of the boundary layer was about 15 mm. (See reference 12 for detailed information on the RJTF.) Table 1 Nominal conditions of RJTF Simulated flight conditions

資料番号: PA1110096000