Supersonic flow fields around Two-Stage-To-Orbit (TSTO) models with different configurations have been experimentally examined in this paper. Four configurations for the orbiter have been considered: A) a hemisphere-cylinder, B) a hemisphere-cylinder with a flat bottom, C) an obliquely truncated circular cylinder, and D) a cone-cylinder. All the flow fields around these models showed complicated shock/shock and shock/boundary-layer interactions, which can be categorized into three patterns, depending on the extent to which the separation shock wave contributes to these interactions. The models B, C and D were proposed to suppress the pressure rise due to the interactions observed in the model A. As a result, the model B showed almost the same interactions as the model A, while in the model C they did not present. In the model D, a large pressure rise was seen in the case with no clearance, whereas the model undergoes the least aerodynamic interaction at a rather large clearance. It is concluded from these results that the model C is less affected by aerodynamic interactions due to the clearance than the other models.
A series of lab-scale firing tests was conducted to investigate the fuel regression characteristics of Cascaded Multistage Impinging-jet (CAMUI) type hybrid rocket. The alternative fuel grain used in this rocket consists of a number of cylindrical fuel blocks with two ports, which were aligned along the axis of the combustion chamber with a small gap. The ports are aligned staggered with respect to ones of neighboring blocks so that the combustion gas flow impinges on the forward-end surface of each block. In this fuel grain, forward-end surfaces, back-end surfaces and ports of fuel blocks contribute as burning surfaces. Polyethylene and LOX were used as a propellant, and the tests were conducted at the chamber pressure of 0.5–2MPa and the mass flux of 50–200kg/m2s. Main results obtained in this study are in the followings: The regression rate of each surface was obtained as a function of the propellant mass flux and local equivalent ratio of the combustion gas. At back-end surfaces the regression rate has a high sensitivity on the gap height of neighboring fuel blocks. These fuel regression characteristics will contribute as fundamental data to improve the optimum design of the fuel grain.
The present paper experimentally investigated boundary-layer transition related with Trailing-Edge Noise (TE noise) emanated from a 2-D NACA0012 airfoil with and without a 2-D thin-tape roughness, which was placed in an anechoic wind tunnel. Surface pressure fluctuation and acoustic measurements showed that the roughness width was sensitive to TE noise generation and magnified the tonal noise when its width corresponded to one and half wavelengths of instability wave in transitional boundary layer. Furthermore, the roughness position had remarkable effects on TE noise generation.
This paper introduces a numerical method to search the steady spin regions for aircraft. A minimization problem is set to calculate the steady spin regions of an aircraft, along with a proposed simple performance index. The spin characteristics are discussed using proposed control surfaces’ margin criteria and their contour maps. Using the YF-16 data, a series of simulations were conducted to check the stable spin states calculated with the method, and control surfaces’ effectiveness around the stable spin states. The simulations’ results show that the proposed contour maps give useful handling information so as to change the steady spin states effectively.
An experimental investigation of a shock-wave/transitional-boundary-layer interaction was conducted on a flat plate with an aft-mounted ramp in a hypersonic gun tunnel. Test Mach number was 10, the Reynolds number range ReL based on the model plate length was from 1.1 to 3.2 × 105, the attack angle of the model was zero, and the ramp deflection angle θw was varied from 10 to 30deg in increments of 2. The effects of ReL and θw on the characteristics of the interaction region were investigated through both schlieren visualization and heat transfer measurement. The state of reattaching boundary layers were determined from these data. The following is obtained. A model that describes how the upstream influence of a shock-wave/boundary-layer interaction scales with the parameters of the incoming flow agrees with laminar data at small ReL-1/8Δθw, where Δθw is the incremental angle after incipient separation. The slope of the characteristic curve of the upstream influence is dependent on the type of the interaction. ReL-1/8 Δθw where the deviation from the model occurs and flow unsteadiness appears is found to become small with increasing ReL.
The dependency of thrust performance on thruster configurations such as antenna length, antenna height, number of antenna, magnetic field configuration, and microwave frequency, was investigated with the objective of improving the thrust performance of microwave discharge ion thruster using antennas for uniform and dense plasma production. The experimental results showed that there was an optimum length of the antennas, and it was 3/4 times the wave length of incident microwaves. The ion beam current reaches its maximum value when the antenna was set at 2mm downstream of an electron cyclotron resonance layer. There was an optimum number of the antennas. This is due to the tradeoff between the coupling of plasma with microwave and the surface recombination on the antenna. The expansion of ionization zone was made successfully by changing magnetic field configuration. In addition, the thrust performance was slightly improved with increase in incident microwave frequency from 2.45GHz to 4.2GHz. A value for the ion beam current with 2.45GHz is compensated by high electron number density and less magnetized ions for the disadvantage of small plasma number density. Overall, the propellant utilization efficiency, ion production cost, and estimated thrust were found to be 0.62, 300W/A and 6.2mN, respectively at mass flow rate of 0.22mg/s for xenon, ion beam voltage of 1,500V and 2.45GHz microwave incident power at 32W.
It is necessary to assess flow qualities more precisely to estimate uncertainties of the wind tunnel test data at higher angles of attack and lower Mach numbers. Then the uncertainties caused by the flow qualities attempted to be identified by repeat tests at different positions in a supersonic wind tunnel. First, the applied conditions were described through the discussions of typical error contributors in the aerodynamic force testing. Next, the standard errors of estimate of the aerodynamic coefficients against angles of attack were calculated from the scattered repeat test data to compare with the uncertainties estimated by the flow calibration data. Finally, the correction factors were introduced to error sensitivities of the flow uniformities to identify the standard errors of estimate caused by the different flow fields. The identified flow quality errors can be applied to estimate data uncertainties of any aerodynamic phenomenon around the similar model shape.
For re-entry missions, deployable-flexible aeroshell is expected to have an advantage because it can significantly reduce the aerodynamic heating due to deceleration at relatively high altitudes thanks to its small mass and large area (i.e. low ballistic coefficient), while it can be stowed compactly during the launch phase. In this study, we propose deployable aeroshell supported by the shape memory alloy frame, which can deploy itself automatically by receiving the aerodynamic heating. The present system neither requires additional sensors nor acuators. It seems suitable for small re-entry capsule because of its simple, fault-tolerant and low-cost nature. It is experimentally demonstrated by the wind tunnel tests that a model having a membrane aeroshell with a shape memory alloy frame is smoothly deployed and kept stable in hypersonic flow.