Yasushi Watanabe

Numerical Analysis on Shock Oscillation of Two-Dimensional External Compression Intakes

Numerical simulations of the unsteady supersonic flow through two-dimensional external compression intakes were performed. The simulations with Euler equations reached to a persistently oscillatory state in subcritical range below some value of the mass flow rate. The frequency of the simulated flow increased with the decrease in the mass flow rate. The intake flows with different inlet Mach numbers, ramp angles and subsonic diffuser area ratios were simulated to examine the effect of those parameters on both of the criteria and frequency of the shock oscillation. The simulations with NavierStokes equations using a low Reynolds number kepsilon turbulence model were also performed, the results of which were compared to the experimental data recently obtained in a supersonic wind tunnel. The shock oscillation frequency predicted by the numerical simulation agrees well with the experimentally obtained frequency, both of which increased with the decrease in the mass flow rate.

EFFECT OF SIDEWALL CONFIGURATIONS ON THE AERODYNAMIC PERFORMANCE OF SUPERSONIC AIR-INTAKE

The effects of the sidewall variation on the aerodynamic performance for a two dimensional external compression supersonic air-intake were investigated by performing both of wind tunnel tests and numerical simulations. It became clear that one of the major disadvantages of the air-intake with “a larger sidewall” is its comparatively poor pressure recovery and skewed spatial distortion, both of which being caused by the separation vortices induced by the interaction between the sidewall boundary layer and the shock waves. It also became clear that another disadvantage is its comparatively large spillage drag. On the other hand, it turned out to have the advantage of comparatively wide stable range in subcritical operation. The reason for the wide stable range of the air-intake with “a larger sidewall” was investigated in detail using both of the results of the wind tunnel tests and numerical simulations.

Computational Simulation of Shock Oscillation Around a Supersonic Air-Intake

Unsteady flow around an external-compression air intake was investigated numerically. Our attention was focused to shock oscillation, which is known as ‘buzz’, in subcritical operation. We used a two-dimensional model consisting of a wedge and a subsonic diffuser. A slit was opened for natural bleed at the entrance of the subsonic duct. The wedge angle was 10 degrees. The free stream Mach number was 1.64. Compressible viscous flow around the model was calculated using UPACS (Unified Platform for Aerospace Computational Simulation) developed at JAXA (Japan Aerospace Exploration Agency). We successfully simulated a series of flow pattern (supercritical, critical, and subcritical operations) by changing the intake mass flow rate. In the subcritical operation, multiple dominant frequencies were found in the pressure fluctuation with the buzz. Low-frequency component was caused by the acoustic resonance of the subsonic diffuser. On the other hand, highfrequency component was generated by the periodic flow separation on the wedge. This phenomenon was caused by the periodic suction of the separated air through the slit. The high-frequency component disappeared when the diffuser length was long. The natural bleed slightly suppressed the pressure fluctuation in the subsonic diffuser in a constant mass flow rate through the diffuser.

LOW BOOM / LOW DRAG SMALL SIZE SUPERSONIC AIRCRAFT DESIGN

Abstract. The French Aerospace Lab (ONERA) and Japan Aerospace Exploration Agency (JAXA) conducted a collaborative research on low-boom and low-drag design of small size supersonic aircraft including engine integration perspective. In this collaborative research, the validation of low-boom design tools, the optimization study on engine position considering aerodynamics, acoustic, and propulsion performances, and the low-boom design of propulsion-airframe integrated (PAI) configuration were performed. In this paper, obtained results are summarized and the low-boom design of PAI configuration is focused on, because the sonic boom loudness is raised by the engine integration. The applied low-boom design tool is based on the free-form deformation and the equivalent area with consideration of the change of lift distribution due to deformations. Results show that the sonic boom loudness of PAI configuration can be reduced to almost the same loudness of glider configuration without engine by the horizontal tail optimization.

Noise Reduction Concept Using Variable Exhaust Nozzle for Supersonic Aircraft

This paper proposes a new variable nozzle for reducing jet mixing noise in supersonic aircraft. The variation of the nozzle is used to achieve a geometry that enhances mixing in the mixing layer without the use of additional devices. Validation tests were conducted using scale models and the thrust loss was evaluated by numerical analysis. It was found that the proposed concept produced an acoustic benefit of 1–2 dB in the overall sound pressure level and a 2.5%–3% thrust loss at takeoff. No cruise thrust penalty is expected considering the clean convergent-divergent configuration of the nozzle under cruise conditions. It was also observed that the characteristics of the nozzle were similar to those of a chevron nozzle, based on which a similar noise reduction performance is expected.

Low Drag Design and Aerodynamic Performance Evaluation of Supersonic Air Inlet

A supersonic air-inlet with bypass technique intended to reduce spillage drag was designed based on a characteristic of an actual jet engine. The effects of bypass design on drag reduction and pressure recovery was examined using CFD simulation. It was found that the bypass design gives a significant drag reduction as well as improves pressure recovery of inlet. The decrement of inlet drag largely surpasses the increment of the nacelle drag within the engine operating range. Furthermore, the bypass duct should be closed when an inlet operates in supercritical condition. Additionally, aerodynamic performance of inlet with diverterless configuration was evaluated to investigate the effects of boundary layer ingestion from airframe surface. In order to improve the inlet performance, a half-conical ramp was adopted to external compression inlet design. It became clear that, thick boundary-layer ingestion causes a huge flow separation within a conventional type inlet, and both pressure recovery and spatial distortion decay drastically. The half-conical ramp prevents boundary-layer from separating, and it is a promising option for diverterless inlet design.

Shock Wave Oscillation Phenomena Depending on Boundary Layer Conditions in Transonic Flow

Shock wave oscillations occur in the internal flow passage. Shock wave oscillations relate with the fluctuation of the boundary layer. The peak frequency of the shock wave oscillations was obtained from the experiment conducted in the cases with the different boundary layer thickness. The results show a change in the peak frequencies depending on the boundary layer conditions. At the high Mach number range, these frequencies have a different trend from previously estimated correlation. Taking into account of the boundary layer conditions and the total pressure losses, we modified the estimation method. The peak frequency by the modified estimation indicates the same trend as that of the experiment.