The drag coefficient for inflatable reentry vehicles shows discrepancies between a wind tunnel experiment and a flight test in a transonic regime. These discrepancies exist in the drag decrease behavior in the transonic regime and local minimum values of drag above the sonic speed in a wind tunnel. Hence, the present paper focuses on uncovering the reasons and mechanisms behind the same and investigates transonic flow fields around a vehicle by using a transonic wind tunnel and the computational fluid dynamics (CFD) approach. Several test models with diameters ranging from 56 to 96 mm are used to quantitatively evaluate the effects of scale and a sting attached on the rear. Aerodynamic coefficients, pressures at the rear of the model, and density gradient distributions are measured for operation conditions of free-stream Mach numbers ranging between 0.8 and 1.3. In addition, detailed distributions of the flow field properties are clarified using the CFD method, which is validated by the experimental data. The results indicate that a sting behind the test models reduces the steep drag decrease at transonic speeds and that shock waves reflected on the test-section walls of the wind tunnel result in local minimum values at supersonic speeds.
Aerodynamic heating around an inflatable reentry vehicle was investigated using hypersonic wind tunnel and numerical approach. The inflatable reentry vehicle is mainly composed of the capsule, membrane aeroshell (rigid in the wind tunnel), and inflatable torus. Basic configuration of the reentry vehicle (HWT-MAAC) is a scaled-down model of SMAAC, which was used in demonstration mission with JAXA / ISAS S-310-41 sounding rocket. Spherical cap of the SMAAC model was replaced by blunt top. Freestream condition of Mach number of 10, reservoir pressure of 2.5 MPa, and reservoir temperature of 950 K was used in the hypersonic wind tunnel test. Heat flux distribution on the surface and density gradient around the HWT-MAAC were measured by infrared thermography and Schliren photograph techniques, respectively. It was found that heat flux distribution widely varies according to angle of attack of the vehicle and a recirculation region near the membrane aeroshell section of the vehicle can appear at high angle of attack. Flow field was also numerically simulated with computational fluid dynamics approach. Analysis solver used here in was RG-FaSTAR, which is a version of JAXA fast aerodynamic routine (FaSTAR). Structures of shock layer and expansion region around HWT-MAAC was discussed through the analysis approach and the wind tunnel results.
scite is a Brooklyn-based organization that helps researchers better discover and understand research articles through Smart Citations–citations that display the context of the citation and describe whether the article provides supporting or contrasting evidence. scite is used by students and researchers from around the world and is funded in part by the National Science Foundation and the National Institute on Drug Abuse of the National Institutes of Health.
hi@scite.ai
10624 S. Eastern Ave., Ste. A-614
Henderson, NV 89052, USA
Copyright © 2024 scite LLC. All rights reserved.
Made with 💙 for researchers
Part of the Research Solutions Family.