This thesis is focused on design control system for Tailless Unmanned Combat Aerial Vehicles. The thesis is divided into two parts: one is a modeling and controller design and verification to stabilize the vehicle even at high angle of attack region and the other is design new concept control surface which delay flow separation and aerodynamic stalling. Thereby improve the effectiveness of wing and other control surfaces. Wind tunnel test was conducted and verified effectiveness of moving vortex generator. Using moving vortex generator, we conduct flight test and verifies stability and better performance.
Unmanned Combat Aerial Vehicles are increasingly and important capability available to military commanders. As they become more capable, they will be replace cruise missiles in many of their traditional roles. As in common with many other modern aircraft, a UCAV with a stealthy design is likely to be aerodynamically unstable. For UCAV to be effective, it will require long endurance and low observability.
The UCAV1303 configuration is a representative model of UCAV. UCAV1303 has no tail wing and has large sweepback angle. For this design UCAV1303 has high nonlinear aerodynamic characteristics, wing rock and pitch break phenomenon. The pitch break makes the aircraft unstable in high angle of attack resign. In high angle of attack resign, it is shown the limit cycle of oscillation in roll mode. It is caused by asymmetric vortices and vortex breakdown.
A wind tunnel test was carried out at the first stage of this research to investigate the aerodynamic characteristics of UCAV1303 with active flow control using micro vortex generator. Wind tunnel test result, MVG delays flow separation so pitch break phenomenon delayed 5°. And micro vortex generator is applied to flight test and verifies better performance and stability.
To control the pitch break and wing rock at high angle of attack region, in this research, L1 adaptive control was implemented with baseline Proportional-derivative(PD) control. L1 adaptive controller is applied to nonlinear simulation and the result illustrate that the system is BIBO stable and non-minimum phase, the L1 adaptive control can find solution to the high angle of attack problem. The L1 adaptive controller can track a desired attitude despite uncertainties and unexpected variation. The performance of the L1 adaptive controller has been validated by numerical simulations and flight test.
미래의 무인전투기는 그 임무에 있어, 국방에서 중요한 역할을 감당하리라 기대된다. 오늘날 전장에서는 적은 피 탐지성을 위해RCS (Radar Cross Section)를 줄이기 위한 큰 뒤젖힘각과 효율적인 비행을 위한 BWB(Blended Wing Body) 형상이 요구되며 그에 따라 UCAV130과 같은 항공기기가 미래 무인투기의 전형적인 대표 형상으로 제시된다.
그러나 이런 비행체는 큰 뒤젖힘각 때문에 앞전 와류에 의한 비선형적 공력 특성을 갖게 되며 이중 pitch break은 비행체의 안정성에 심각한 문제를 일으킨다. 따라서 본 연구에서는 UCAV1303을 중심으로 이런 기체의 안정성 증대를 위한 장치로 가동형 윙펜스를 설계하고 비행실험에 적용하였으며 L1 적응 제어기를 설계하고 비행실험을 통해 제어기의 성능을 검증하였다.
풍동실험을 통해 Micro Vortex Generator가 앞전 와류를 막아 Pitch Break 현상을 지연 시키는 결과를 얻었다. 이를 비행실험에 적용하기 위하여 MVG의 크기를 키워 Wing Fence형태로 장착하여 앞전 와류를 막아주고 안정적인 비행실험을 수행하였다. 레이더 탐지 면적에 영향을 주지 않기 위해 Wing Fence를 가동형으로 설계하여 안정적인 이착륙 및 비행실험을 수행하였다.
피치 각 제어를 위하여 baseline 제어기로 PD제어기를 설계하였고