Search

The series of aerodynamic heating test campaigns with blunt shaped models in hypersonic high-enthalpy flow were conducted in JAXA-HIEST. Four different blunt shaped models were applied in this study; Apollo CM 6.4% scale model, HRV capsule 6% scaled model and two hemi-sphere probes with radius of curvature r=20mm and r=100mm. Numerical analysis with an axis-symmetric Navier-Stokes thermo-chemical non-equilibrium code was also conducted. Stagnation pressure was almost fixed to P0=50MPa, while stagnation enthalpy was varied from H0=7MJ/kg to 21MJ/kg. The unit Reynolds number under the condition was 0.9million/m to 3million/m and the boundary layer was hence mainly fully laminar flow. Abnormal high-heat flux was observed under high enthalpy condition (H0>10MJ/kg), which anomalies were almost twice higher than those of numerical calculations or theoretical ones. The measurement with four different blunt models showed that the abnormal heat flux was increased as stagnation enthalpy. From the analysis of compression process of the free-piston driver, radiation heating from driver gas does not seem to explain this heating anomaly. The numerical analysis implied that the heating anomaly was related to radiation from unknown atoms or molecules in shock layer . Senior Researcher, Space Transportation Propulsion Research and Development Center, Space transportation mission directorates, tanno.hideyuki@jaxa.jp, Member AIAA Senior Researcher, Space Transportation Propulsion Research and Development Center, Space transportation mission directorates Deputy Manager, Space Transportation Propulsion Research and Development Center, Space transportation mission directorates, Senior member AIAA Graduate student, Department of Aerospace Engineering, ishihara@cfd.mech.tohoku.ac.jp, member AIAA Assistance Professor, Department of Aerospace Engineering, yogi@cfd.mech.tohoku.ac.jp, Member AIAA Professor, Department of Aerospace Engineering, sawada@cfd.mech.tohoku.ac.jp, AIAA Associate Fellow 43rd AIAA Thermophysics Conference 25 28 June 2012, New Orleans, Louisiana AIAA 2012-3104 Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. American Institute of Aeronautics and Astronautics 2

Surface pressure fluctuation was measured in hypersonic high-enthalpy transitional boundary layer. The measurement mainly focused on observation of the second mode instability in the transition process under hypersonic high-enthalpy flow with high-speed pressure transducers. A 7-degree half-angle cone model with 1100mm length was newly manufactured for this study and a wind tunnel test campaign with the model was conducted in the high-enthalpy shock tunnel JAXA-HIEST. On the surface of the model twenty-six pressure transducers (PCB132) were mounted axially from x=412mm to x=1052mm (x means distance from the cone apex) approximately 40mm interval to observe frequency spectrum of the pressure fluctuation in boundary layer. In addition, eighty-eight miniature co-axial thermocouples were mounted to determine boundary layer status and transition location observing heat flux distributions. In the present test campaign stagnation enthalpy was varied from H0=2.5MJ/kg to 11MJ/kg, while free-stream Reynolds number will be hold to constant value (Unit Reynolds number is approximately 2x10 6 to 3x10 6 /m) with increasing stagnation pressure from 13MPa (H0=3.5MJ/kg) to 60MPa (H0=9.4MJ/kg) as enthalpy rise. To detect the high temperature gas effects including oxygen dissociation and vibration excitation, air and pure nitrogen test gas were applied in this test campaign. The remarkable difference between air and pure nitrogen test gas has not detected in the campaign.

A miniature data-logger was developed for aerodynamic measurement in the free-piston shock tunnel HIEST. The data-logger was designed to be mounted on wind tunnel test models, with an overall size of 50 mm x 70 mm x 40 mm, including batteries. Since the logger was designed to use piezoelectric sensors (ICP type), it includes four piezoelectric amplifiers and four power supplies, as well as high-speed analog-digital converters, which digitize the measured data with 16-bit resolution. The logger’s sampling rate and sample size are 500 kHz and 200 ms, respectively. The logger waits for a trigger signal (accelerometer output) and then starts to take measurements. The trigger threshold and pre-trigger delay time can be adjusted arbitrarily. Measured data is stored to static memories for transfer to a PC via a USB interface after a wind tunnel test. To demonstrate the entire data-logger system, four wind tunnel shots were conducted in HIEST. A cone model with blunted nose (nose radius 20 mm) was used with piezoelectric accelerometers (PCB) 352C66 and a piezoelectric pressure transducer (PCB 113A21) inside the model. In the present wind tunnel test campaign, records of pressure and axial force were obtained under conditions of H0 = 4 MJ/kg, P0 = 14 MPa and H0 = 7 MJ/kg, P0 = 17 MPa. This demonstrated that the system worked correctly in the short test duration and harsh conditions typical of HIEST. Use of this data-logger allows the elimination of a large-diameter sting, ending concerns about the sting’s interference with the base flow of the model, which could cause serious errors in measurement in wind tunnel tests.

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 0o and 30o. Stagnation enthalpy and stagnation pressure were varied from H0=4MJ/kg to 21MJ/kg, and from P0=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 H0=4MJ/kg to 12MJ/kg, implying that the boundary layer flow was usually laminar. However, shots under high enthalpy and high Reynolds number conditions (H0 over 18MJ/kg and P0 approximately 50MPa) showed that the measured heat flux was approximately 40% higher than that obtained in other conditions.

A feasibility study of a force measurement technique for impulsive facilities was performed in the free-piston high enthalpy shock tunnel HIEST. The technique is simple and cost effective rather than conventional aerodynamic force balance technique. With the present technique, aerodynamic force was obtained as products of measured acceleration and mass of the model, which model was weakly suspended with thin wires. Natural vibration of the test model, which often degrade measurement accuracy and time response, can be removed through signal recovery process with a de-convolution calculation and with low-path filtering. Time response of the measurement technique guaranteed 0.1ms with a digital low-path filter, which cut-off frequency is 10kHz. For evaluation of the technique, wind tunnel test in HIEST was conducted to measure aerodynamic force of a 500mm length HB-2 standard model. The wind tunnel test results showed that the present technique can guarantee time response as order of sub-ms without high-frequency messy noise. The comparison was also performed between the present results and the results obtained in the hypersonic blow-down wind tunnel HWT2.