• Advances in aircraft and spacecraft science
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• 제2권2호
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• pp.169-182
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• 2015

Aeronautics engine cooling is one of the biggest problems that engineers have tried to solve since the beginning of human flight. Systems like radiators should solve this purpose and they have been studied extensively and various solutions have been found to aid the heat dissipation in the engine zone. Special interest has been given to air coolers in order to guide the air flow on engine and lower the high temperatures achieved by the engine in flow conditions. The aircraft companies need faster and faster tools to design their solutions so the development of tools that allow to quickly assess the effectiveness of an cooling system is appreciated. This paper tries to develop a methodology capable of providing such support to companies by means of some application examples. In this work the development of a new methodology for the analysis and the design of oil cooling systems for aerospace applications is presented. The aim is to speed up the simulation of the oil cooling devices in different operative conditions in order to establish the effectiveness and the critical aspects of these devices. Steady turbulent flow simulations are carried out considering the air as ideal-gas with a constant-averaged specific heat. The heat exchanger is simulated using porous media models. The numerical model is first tested on Piaggio P180 considering the pressure losses and temperature increases within the heat exchanger in the several operative data available for this device. In particular, thermal power transferred to cooling air is assumed equal to that nominal of real heat exchanger and the pressure losses are reproduced setting the viscous and internal resistance coefficients of the porous media numerical model. To account for turbulence, the k-${\omega}$ SST model is considered with Low- Re correction enabled. Some applications are then shown for this methodology while final results are shown in terms of pressure, temperature contours and streamlines.

• Advances in aircraft and spacecraft science
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• 제9권5호
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• pp.415-431
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• 2022

Secondary flows have a huge impact on losses generation in modern low pressure gas turbines (LPTs). At design point, the interaction of the blade profile with the end-wall boundary layer is responsible for up to 40% of total losses. Therefore, predicting accurately the end-wall flow field in a LPT is extremely important in the industrial design phase. Since the inlet boundary layer profile is one of the factors which most affects the evolution of secondary flows, the first main objective of the present work is to investigate the impact of two different inlet conditions on the end-wall flow field of the T106A, a well known LPT cascade. The first condition, labeled in the paper as C1, is represented by uniform conditions at the inlet plane and the second, C2, by a flow characterized by a defined inlet boundary layer profile. The code used for the simulations is based on the Discontinuous Galerkin (DG) formulation and solves the Reynolds-averaged Navier-Stokes (RANS) equations coupled with the Spalart Allmaras turbulence model. Secondly, this work aims at estimating the influence of viscosity and turbulence on the T106A end-wall flow field. In order to do so, RANS results are compared with those obtained from an inviscid simulation with a prescribed inlet total pressure profile, which mimics a boundary layer. A comparison between C1 and C2 results highlights an influence of secondary flows on the flow field up to a significant distance from the end-wall. In particular, the C2 end-wall flow field appears to be characterized by greater over turning and under turning angles and higher total pressure losses. Furthermore, the C2 simulated flow field shows good agreement with experimental and numerical data available in literature. The C2 and inviscid Euler computed flow fields, although globally comparable, present evident differences. The cascade passage simulated with inviscid flow is mainly dominated by a single large and homogeneous vortex structure, less stretched in the spanwise direction and closer to the end-wall than vortical structures computed by compressible flow simulation. It is reasonable, then, asserting that for the chosen test case a great part of the secondary flows details is strongly dependent on viscous phenomena and turbulence.

• 한국항공우주학회지
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• 제43권2호
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• pp.109-117
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• 2015

덕트는 항공기의 내부엔진에 외부 공기를 흡입하기 위한 장치이다. 엔진 입구면의 레이더 반사량을 줄여 피탐지성을 감소시키기 위하여 S형태의 덕트를 가지게 되었다. S-Duct는 중심선의 곡률, 입구형상 등의 형상변수에 따라 엔진의 성능에 영향을 미친다. 본 연구에서는 RAE M 2129 S-Duct의 입구형상에 대하여 가로세로비의 변경에 따른 덕트 내부 유동에 대한 유동 특성을 알아보기 위해 전산해석을 수행하였다. S-Duct의 성능 평가 기준으로는 유동 왜곡계수를 사용하였다. 공력해석을 위해 상용해석 소프트웨어를 사용하였으며, 벽면에서의 역압력 구배의 영향으로 발생하는 유동박리와 2차 유동을 예측하기 위하여 $k-{\omega}SST$ 난류모델을 사용하였다. S-Duct의 Port side와 Starboard side 각각의 압력분포 값에 대하여 ARA의 실험값과 비교하여 본 연구에서 사용된 전산해석 기법의 타당성을 검증하였다. 해석 결과 모든 형상에 대하여 유동박리와 2차 유동이 발생하는 것을 확인하였다. 반원형 형태의 입구형상을 가지는 S-Duct가 뛰어난 성능을 보임을 확인하였다.

• Advances in aircraft and spacecraft science
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• 제3권4호
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• pp.367-378
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• 2016

Thrust Vectoring is a dynamic feature that offers many benefits in terms of maneuverability and control effectiveness. Thrust vectoring capabilities make the satisfaction of take-off and landing requirements easier. Moreover, it can be a valuable control effector at low dynamic pressures, where traditional aerodynamic controls are less effective. A numerical investigation of Fluidic Thrust Vectoring (FTV) is completed to evaluate the use of fluidic injection to manipulate flow separation and cause thrust vectoring of the primary jet thrust. The methodology presented is general and can be used to study different techniques of fluidic thrust vectoring like shock-vector control, sonic-plane skewing and counterflow methods. For validation purposes the method will focus on the dual-throat nozzle concept. Internal nozzle performances and thrust vector angles were computed for several range of nozzle pressure ratios and fluidic injection flow rate. The numerical results obtained are compared with the analogues experimental data reported in the scientific literature. The model is integrated using a finite volume discretization of the compressible URANS equations coupled with a Spalart-Allmaras turbulence model. Second order accuracy in space and time is achieved using an ENO scheme.

• 제어로봇시스템학회논문지
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• 제14권1호
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• pp.95-103
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• 2008

This paper deals with the problem of autolanding for aircraft under windshear environment for which the landing trajectory is given. It is well known that the landing maneuver in windshear turbulence is very dangerous and hard for the pilot to control because windshear is unpredictable in when and where it happens and its aerodynamic characteristics are complicated. In order to accomplish satisfactory autolanding maneuver in this environment, we propose a gain-scheduled controller. The proposed controller consists of three parts: PID controller, called baseline controller, which is designed to satisfy requirements of stability and performance without considering windshear, gain scheduler based on fuzzy logic, and safety decision logic, which decides if the current autolanding maneuver needs to be aborted or not. The controller is applied to a 6-DOF simulation model of the associated airplane in order to illustrate the effectiveness of the proposed control algorithm. It is noted that a cross wind in the lateral direction is included to the simulation model. From the simulation results it is observed that the proposed gain scheduled controller shows superior performance than the case of controller without gain scheduling even in severe downburst and tailwind region of windshear. In addition, touchdown along centerline of the runway is more precise for the proposed controller than for the controller without gain scheduling in the cross wind and the tailwind.

• 한국생산제조학회지
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• 제20권6호
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• pp.714-719
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• 2011

The surface roughness depending on the machining method is very important because is produce a finished product through riveting, sealing, bonding, and special paint in order to curb the turbulence and air resistance which occur between the sheets. Aluminum alloy 2024 which is widely used for interior and exterior material of aircraft are tested. Jin-young JW-60C wire cutting machine was used in this experiment. In this paper, the experimental investigation has been performed to find out the influence of the surface roughness and surface shape characteristics on the wire-cut EDM of discharge energy in aluminium alloy 2024. The selected experimental parameters are peak current, no-load voltage, off time and feed rate. The experimental results give the guideline for selecting reasonable machining parameters. The high discharge energy on the idle time, almost no change in surface roughness can be seen.

• 한국추진공학회지
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• 제23권4호
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• pp.10-20
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• 2019

충격파 개념을 이용하는 유체 추력벡터 제어는 빠른 벡터링 응답, 간단한 구조 및 낮은 무게로 인하여 큰 벡터링 성능을 달성하는데 많은 이점을 제공한다. 본 논문에서는 전산유체역학 기법을 사용하여 슬롯 인젝터를 가진 3차원 직사각형 초음속 노즐에 대하여 연구를 수행하였다. 계산 방법론을 검증하기 위하여 수치 결과를 실험 데이터로 비교하였다. 대칭 평면에서의 상부 및 하부 노즐벽을 따르는 압력분포는 시험 결과와 잘 일치하였다. $k-{\omega}$ SST 난류모델을 기반으로 한 수치해석을 통하여, 운동량 플럭스 비율의 영향을 철저히 조사하여 추력의 성능 변화를 명확하게 나타내었다.

• Advances in aircraft and spacecraft science
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• 제10권3호
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• pp.203-222
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• 2023

The detonation combustion is a supersonic combustion process follows on shock wave oscillations in detonation tube. In this paper numerical studies are carried out combined effect of blockage ratio and spacing of obstacle on detonation wave propagation of hydrogen-air mixture in pulse detonation combustor. The deflagration to detonation transition of stoichiometric (ϕ=1)fuel-air mixture in channel has been analyzed for effect of blockage ratio (BR)=0.39, 0.51, 0.59, 0.71 with spacing of 2D and 3D. The reactive Navier-Stokes equation is used to solve the detonation wave propagation mechanism in Ansys Fluent platform. The result shows that fully developed detonation wave initiation regime is observed near smaller vortex generator ratio of BR=0.39 inside the combustor. The turbulent rate of reaction has also a great significance role for shock wave structure. However, vortices of rapid detonation wave are appears near thin boundary layer of each obstacle. Finally, detonation combustor demonstrates the superiority of pressure gain combustor with turbulent rate of reaction of 0.6 kg mol/m3 -s inside the detonation tube with obstacle spacing of 12 cm, this blockage enhanced the turbulence intensity and propulsive thrust. The successful detonation wave propagation speed is achieved in shortest possible time of 0.031s with a significance magnitude of 2349 m/s, which is higher than Chapman-Jouguet (C-J) velocity of 1848 m/s. Furthermore, stronger propulsive thrust force of 36.82 N is generated in pulse time of 0.031s.

• 한국항공우주학회지
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• 제50권12호
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• pp.899-907
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• 2022

초경량 비행장치인 드론의 최대 허용 비행 고도는 지상 150m로 이는 난류의 영향을 받아 바람의 변동성이 강한 대기경계층 내에 존재한다. 또한 대기경계층 내에서의 바람 변동성은 지리적 위치에 따라 다른 특성을 가지므로 드론 관련 안전사고 방지를 위해서는 비행 지역에서의 각 고도의 바람 특성에 대한 명확한 이해가 필요하다. 본 연구에서는 인천국제공항 인근에 위치한 항공기상관측장비 테스트베드에서 윈드라이다(WindMast 350M)를 사용하여 2022년 7월과 9월에 바람의 연직 구조 관측을 수행하였고, 이러한 관측된 바람 자료를 활용하여 드론의 안전비행을 위한 정보를 생산하는 분석 방안을 제시하였다. 우선 윈드라이다를 통해 수집된 바람 자료에 푸리에 변환 분석 방법을 사용하여 수평 풍속의 시간 규모 특징을 각 고도별로 살펴보았다. 또한 강수와 무강수 사례의 바람장의 스펙트럼으로부터 드론 비행에 중요한 바람의 시간 규모인 1시간 이하 규모의 수평 풍속의 분산을 분리하여 전체 규모에 대한 1시간 이하 규모의 기여도를 각 고도별로 확인하였다.

• 한국항공우주학회지
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• 제42권2호
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• pp.99-107
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• 2014

본 연구에서는 원격 조종 소형 비행기에서 주로 사용되고 있는 Kline-Fogleman 익형의 저 레이놀즈수 공력 특성을 분석하는 연구를 수행하였다. NACA4415와 이를 기반으로 한 Kline-Fogleman 익형의 공력특성을 비교하였다. 본 연구는 ANSYS Fluent를 활용하였으며, 유동은 비압축성으로 가정하고, 난류모델 $k-{\omega}$ SST를 사용하였다. 이를 통하여 Kline-Fogleman 익형의 공기역학적 원리를 규명하였으며 계산된 레이놀즈수 $3{\times}10^3{\sim}3{\times}10^6$ 범위에서 Kline-Fogleman 익형이 NACA4415에 비해 양력계수가 향상됨을 확인하였다. 특히 레이놀즈수 $2.4{\times}10^5$이하의 영역에서는 Kline-Fogleman 익형의 양항비가 NACA4415에 비해 26%까지 향상되었다.