Yeol Lee

Fluidic Thrust Vector Control of Supersonic Jet Using Co-flow Injection

The purpose of this research is to investigate the operation condition of fluidic thrust vector using injection of the control flow tangential to the main jet direction; co-flow injection. The physical model of concern includes a chamber and a supersonic nozzle for supersonic main jet injection, and two chambers with slots for control flow injection. Steadystate numerical and experimental studies were conducted to investigate operating parameters; detailed flow structures, jet deflection angles, and shock effects were observed near the nozzle exit. An unsteady numerical calculation was conducted to analyze the dynamic characteristics of fluidic control of jet vectoring up- and downward from the nozzle axis, so that the response time of jet deflection to control flow injection and the pressure dispersion on the nozzle wall were investigated. Internal nozzle performance was predicted for total pressure range of the jet from 300 kPa to 1000 kPa to the control flow pressure from 120 to 200 kPa. To take into account the important features of high-speed flows, including shock-boundary layer interactions, a low Reynolds number k-e turbulence model with compressible-dissipation and pressure-dilatation effects was applied.

Development of the High-Accuracy Multi-Component Balance for Fluidic Thrust Vectoring Nozzle of UAV

The thrust vector control technique is essential for high maneuverability of unmanned aerial vehicles. In the present study, a multi-component balance was developed to quantitatively evaluate the thrust-vectoring performance of a supersonic rectangular nozzle based on the Coanda coflowing effect. Precise calibration and detailed data analysis were performed during the development. It was found that the cross-talk errors between load cells in the balance were less than 5%, and that the unwanted errors due to high-pressure supply tubes were almost negligible, which contributed to the high accuracy of the present balance design. Some preliminary test results of the thrust-vectoring performance of the present nozzle design were also obtained and analyzed.

Skin Friction Measurements for Recirculating Normal-Shock/Boundary-Layer Interaction Control

Skin-friction measurements for a normal-shock/boundary-layer interaction with several recirculating flow control methods have been conducted in a planar Mach 1.4 wind tunnel. The skin friction has been measured along the spanwise centerline and downstream of the interaction using the laser interferometry skin-friction (LISF) technique, which optically detects the rate of thinning of an oil film applied to the test surface. Velocity profiles measured by laser Doppler velocimetry at the same locations as the LISF measurements in the interaction were also used to evaluate the skin friction. Comparison of the two measurement techniques for skin friction is found to show reasonable agreement. Various configurations of the mesoflap arrays of different shapes and thicknesses were examined, and the results were compared to cases of both a solid wall with no control mechanisms and conventional recirculating passive flow control with a porous plate. Of the various mesoflap arrays tested, one mesoflap array provided higher skin friction downstream of the interaction, and as such tends to have better recovery from flow separation. However, all values of skin friction for the mesoflap arrays and the porous plate were found to be lower than for the solid-wall reference case. As such, the flow downstream of these control systems can be more susceptible to separation for this particular condition, although the control systems may reduce viscous drag in external flows.

Application of Backstep Coanda Flap for Supersonic Coflowing Fluidic Thrust-Vector Control

Pb = atmospheric pressure Pt1 = primary-nozzle operating total pressure Pt2 = secondary-nozzle operating total pressure Pw = wall static pressure on the coanda flap s = height of the secondary nozzle sh = backstep height of the coanda flap T = resultant thrust, Tx Tz 1∕2 Tx = axial thrust Tz = side thrust t = lip thickness of the primary-nozzle exit δp = resultant pitch thrust-vector angle, tan −1 Tz∕Tx

Micro-vortex generators and recirculating flow control of normal shock stability and position sensitivity

Experiments have been conducted herein at Mach 1.4 to determine the effect on normal shock stability of a micro-ramp array, ramped vane arrays, porous plates over a cavity, and an open cavity as compared to no control (solid wall) in the region of a 5 o diffuser. For each of the control devices, the mean position and standard deviation in position of the normal shock were determined for a set stagnation pressure over the range of tunnel stagnation pressures from 134.4 to 148.2 kPa using schlieren photography. For each of the set stagnation pressures, the boundary layer static pressure fluctuations downstream of the diffuser entrance were also obtained for each control device. An array of three micro-ramps and an array of three ramped vanes, whose heights were scaled to 40% of the incoming boundary layer thickness, were placed ahead of the normal shock, as well as an array of two ramped vanes with a height of 0.6 the incoming boundary layer thickness. Three porous plate configurations with a porosity of 5% were also examined with varying lengths and positions. The open cavity used was 13.5 and 5 incoming boundary layer thicknesses long and deep, respectively. It was demonstrated that the control devices did affect shock stability in the region of the diffuser, both increasing and decreasing shock stability depending on shock position. Schlieren photography showed the micro-vortex generators reduced fluctuations of normal shock position when the shock moved into the diffuser, while slightly increasing fluctuations in shock position when the mean shock position was upstream of the diffuser entrance. The porous plate that terminated 3.5 incoming boundary layer thicknesses upstream of the diffuser entrance reduced the shock position fluctuations near the diffuser entrance. High frequency wall pressure measurements showed that the open cavity should not be used to improve stability. Also, the standard deviation of the wall static pressure fluctuations inside the diffuser was improved with the micro-vortex generators when the mean shock position was near and downstream of the diffuser shoulder.

Parametric Investigation on the Essential Flow Factors Commanding Steady Operations of the Second Throat Exhaust Diffuser

The constant area exhaust diffuser (CAED) demands a higher starting pressure than the second throat exhaust diffuser (STED). In the design process, one dimensional (1D) normal shock theory was explored to obtain the optimum starting pressure of STED. Experiments were conducted with a small scale cold gas simulator using nitrogen gas as injectant. The application of 1D normal shock theory for STED design was examined with 20-25% difference in each accuracy in comparison to numerical and experimental evidences. The relation between evacuation quality and essential geometrical parameters such as diffuser– to–nozzle throat area ratio (Ad/At), diffuser–to–the second throat area ratio (Ad/As), and the nozzle expansion ratio (e) is presented for the optimization of STED. The optimum starting pressure increases in proportion to Ad/At and the optimized Ad/As was predicted in the range of 2.2-2.5. After STED starting, the evacuation quality is the same whether the expansion ratio of the nozzle is large or not. However, the range of the transition regime varied according to the nozzle expansion ratio. Nomenclature d A = area of diffuser inlet s A = area of the second throat t A = area of nozzle throat D = diameter e

Experimental Study of the Effect of Side Plate on the Coanda Effect of Sonic Jet

ABSTRACT An experimental study for the characteristics of the thrust-vectoring of a sonic jet utilizing the coanda flap installed at a rectangular nozzle exit is performed. Two side plates are installed at both sides of the flap to decrease the three dimensional effects of the jet on the flap surface. Schlieren flow visualizations and quantitative measurements of the deflection angle of thrusting vector show that the side plates are able to delay the separation of the jet at the downstream of the flap surface. Substantial increase in the deflection angle of the jet as high as 72˚ and small thrust loss as low as 7% are obtained by the present thrust-vectoring technique using the side plates.초 록 사각노즐에서 발생한 음속제트의 코안다 효과를 이용한 추력편향 제어에 관한 실험적 연구가 수행되었다. 코안다 플랩 표면에서 나타나는 제트유동의 3차원 효과를 저감시키기 위하여 노즐 출구의 코안다 플랩 양쪽에 측판이 설치되었다. 쉴리렌 유동가시화 기법과 정량적인 추력편향각 측정을 통하여, 플랩 양쪽에 설치된 측판에 의하여 제트유동의 플랩 하류에서의 박리현상이 크게 지연되었음을 관찰하였다. 이에 따라 최대 72도의 높은 추력편향각과 약 7% 정도의 적은 추력손실이 관찰되었다.Key Words:Thrust Vector Control(추력편향제어), Sonic Jet(음속제트), Shock Wave(충격파), Coanda Effect(코안다 효과), Flow Separation(유동박리) Received 14 August 2015 / Revised 8 March 2016 / Accepted 15 March 2016

Experimental Study of the Quantitative Characteristics of Fluidic Thrust Vectoring Nozzle for UAV

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Application of Back-Step Coanda Flap for the Supersonic Co-flowing Fluidic Thrust Vector Control

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A Study on Operation Characteristics of Co-flow Fluidic Thrust Vector Control under Over-expanded Jet Condition

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