Free-piston Shock Tunnel HIEST Research Articles

Aerodynamic characteristics of a free-flight scramjet vehicle in shock tunnel

A multi-component aerodynamic test for an airframe-engine integrated scramjet vehicle model was conducted in the free-piston shock tunnel HIEST. A free-flight force measurement technique was applied to the scramjet vehicle model named MoDKI. A new method using multiple piezoelectric accelerometers was developed based on overdetermined system analysis. Its unique features are the following: (1) The accelerometer’s mounting location can be more flexible. (2) The measurement precision is predicted to be improved by increasing the number of accelerometers. (3) The angular acceleration can be obtained with single-axis translational accelerometers instead of gyroscopes. (4) Through the averaging process of the multiple accelerometers, model natural vibration is expected to be mitigated. With eight model-onboard single-axis accelerometers, the three-component aerodynamic coefficients (Drag, Lift, and Pitching moment) of MoDKI were successfully measured at the angle of attack from 0.7 to 3.4 degrees under a Mach 8 free-stream test flow condition. A linear regression fitting revealed a 95% prediction interval as the measurement precision of each aerodynamic coefficient.Graphical abstract

S1910103 高次精度スキームを用いた大気圏突入カプセル後方部の強制遷移流れ場計算([S191-01]大気突入・減速技術(1),宇宙工学部門)

An aeroheating measurement test campaign on an Apollo capsule model with turbulent boundary layers was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. When flow became turbulent flow, heat flux on fore and rear body was 1.5-2 times larger than one in laminar flow. In order to reproduce the heat flux by numerical simulation, we constructed high order CFD code towards LES in hypersonic flow. Since robustness near shock and high resolution in boundary layer is needed for LES in hypersonic flow, we employed recently improved WENO method. We calculated the hypersonic flow (M_∞ = 17) around cylinder and investigated the robustness for strong shock by this method. We could obtain smooth pressure distribution and good agreement with the calculated heat flux by NASA LAURA code. The hypersonic flow field around HTV-R, which is a manned space capsule under development by JAXA was also calculated towards predicting heat flux on after body.

極超音速乱流場解析に向けたWENO法に基づくLESコード開発

An aeroheating measurement test campaign on an HTV-R capsule model with turbulent boundary layers was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. When flow became turbulent flow, heat flux on rear body was 2-3 times larger than one in laminar flow. In order to reproduce the turbulent heat flux by numerical simulation, we constructed high order CFD code towards LES in hypersonic flow. Since robustness near shock and high resolution in boundary layer is needed for LES in hypersonic flow, we employed recently improved WENO method. We calculated the hypersonic flow (M_∞=17) around cylinder and investigated the robustness for strong shock by this method. We could obtain smooth pressure distribution and good agreement with the calculated heat flux by NASA LAURA code. The hypersonic flowfield around HTV-R was also calculated towards predicting turbulent heat flux on after body.

0902 高エンタルピ衝撃風洞での境界層遷移トリップの開発(OS9-1 超音速流れの基礎と応用,OS9 超音速流れの基礎と応用)

Development of boundary layer trip for capsule shaped reentry vehicles under high-enthalpy hypersonic condition was developed in the free-piston shock tunnel HIEST. Wind tunnel test campaign with an 6% scaled Apollo CM model was conducted with stagnation pressure of 30MPa to 60Mpa and with stagnation enthalpy from 3.5MJ/kg to 22MJ/kg. Reynolds number based on model diameter was varied from 0.2 to 1.3 million. To measure heat flux around the model, eighty-four miniature co-axial thermocouples were instrumented on the heat shield surface of the model. Angle-of-attack of the model was fixed to 28degree. A pizza-box configuration boundary layer trip was mounted on the heat shield surface of the model, which trip height k were 0.3mm, 0.6mm and 0.8mm. With k=0.6mm height trip, boundary layer became fully turbulent less than H0=10MJ/kg. However, boundary layer was still laminar over H0=15MJ/kg even with k=0.8mm trip.

Improvement of a Model Scramjet Engine Design Using Various Types of Ground Tests

The scramjet, one of the most promising candidates for future transport systems, has many advantages such as a simple configuration and a high specific impulse. However, various characteristics of scramjet engines in a hypersonic environment are difficult to predict from theory. In the present study, we investigated the characteristics and the performance of the scramjet engine with various types of ground tests. An engine model (S1) was tested in the free-piston shock tunnel T4. (S1) showed active supersonic combustion and good starting characteristics, but its operating range is rather limited. To improve the performance of the model scramjet engine, we also investigated major components, such as the combustor and the intake by experiments with different model configurations. In the combustor test results, the zigzag cavity flame holder showed the best combustion performance among the three cases. By testing various intake configurations, the most efficient intake configuration was also determined. A revised model (S2), which was redesigned based on the component test results, showed enhanced thrust performance and a wider operating range in the free jet test using the free piston shock tunnel HIEST.

S1904-2-1 高圧・高エンタルピ流れにおけるApollo模型の空力過熱(大気突入・減速技術(2),社会変革を技術で廻す機械工学)

Aeroheating of an Apollo CM model was measured in the free-piston shock tunnel HIEST. The primary purpose of the measurement was to obtain benchmark data for numerical codes validation under high-enthalpy high-Reynolds number conditions. The model was made of SUS-304 stainless steel and had a diameter of 250mm. The windward surface of the model was equipped with 84 miniature co-axial Chromel-Constantan thermocouples that were manufactured in-house at JAXA. Twelve thermocouples and two Piezo-resistive pressure transducers were also mounted aft on the model and were used to determine the establishment of flow around the model. In this test campaign, heat flux data were measured for angles of attack 0° and 30°. Stagnation enthalpy and stagnation pressure were varied from H_0=4MJ/kg to 21MJ/kg, and from P_0=14MPa to 58MPa, respectively. Aeroheating characteristics were observed with a fully laminar boundary layer and with a transition boundary layer. Measured heat flux was normalized by the product of the Stanton number and the square-root of the Reynolds number for correlation with other wind tunnel results and flight results. The normalized heat flux value did not change remarkably when stagnation enthalpy was changed from H_0=4MJ/kg to 12MJ/kg, implying that the boundary layer flow was usually laminar. However, shots under high enthalpy and high Reynolds number conditions (H_0 over 18MJ/kg and P_0 approximately 50MPa) showed that the measured heat flux was approximately 40% higher than that obtained in other conditions.

Characteristics of Free-Piston Shock Tunnel HIEST. 1st Report. Tuned Operation of Free-Piston Driver.

A free-piston shock tunnel is the most capable facility to reproduce hypervelocity flows with chemical nonequilibrium around reentry vehicles. However, performance expected for the free-piston shock tunnel has not been achieved yet, mainly due to the lack of driver-gas volume after free-piston compression, which causes decay of shock wave strength and reduction in test time. In order to overcome this problem, the authors proposed the tuned operation method which enables safe overdriving of piston to maintain driver-gas pressure, and ensured its applicability by the tests in the middle-size pilot facility. Then, based on the tuned operation theory the largest free-piston shock tunnel HIEST was designed and built in the National Aerospace Laboratory. In the present study, the tuned operation of the HIEST was attempted and the effect on the performance of the free-piston driver was discussed.