• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.41 no.6
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• pp.488-494
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• 2013

The control law design and analysis environment of the FBW helicopter system have been developed using model base design method. The model based design is generally used in a aircraft, unmanned aerial system and automobile as well as rotorcraft development. The model based design provides many advantages such as development risk and schedule reduction using simulation and autocode generation. This paper describes a development of process for verification and validation of auto generated code for FBW helicopter flight control law. And this process is applied for Fly-By-Wire Helicopter Development Project. The results of functional test for auto generated code meet several specific requirements.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.7
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• pp.85-91
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• 2005

Supersonic jet fighter aircraft must have been guaranteed appropriate for controllability and stability in HAoA(High Angle of Attack) region. Limit value of aircraft enter the deep stall at HAoA is related to problem of aircraft configuration design. But, In order to guarantee the aircraft safety in HAoA, control law is designed using digital Fly-By-Wire flight control system in modern versions of supersonic jet fighter aircraft. Also, In order to recovery if aircraft enter the deep stall or spin, anti-spin control law and MPO(Manual Pitch Override) mode is designed. AoA limiter and MPO is designed in longitudinal axis and HAoA departure prevention logic, roll command limiter, rudder fader and anti-spin logic is designed in lateral-directional axis. In this paper, we introduce the T-50 HAoA flight control law and propose that aircraft stability and adequate of these control law from HAoA flight test.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.5
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• pp.41-47
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• 2002

In this paper, a guidance law applicable to aircraft automatic landing is proposed and its performance is compared with the conventional ILS-type landing approach. The concept of miss distance, which is commonly used in the missile guidance laws, and Lyapunov stability are effectively combined to obtain the landing guidance law. The new landing guidance law is integrated into the existing controller and is applied to the landing approach and flare phases of landing procedure. Numerical simulation results show that the new landing guidance law is a viable alternative to the conventional strategies that directly control the longitudinal deviation or altitude.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.10
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• pp.60-67
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• 2005

Modern version of supersonic jet fighter aircraft must have guaranteed appropriate controllability and stability in HAoA(high angle of attack). Limit value of aircraft entering into the deep stall in HAoA is related to aircraft configuration design. But, In order to guarantee the aircraft's safety in HAoA, control law for HAoA region implemented in digital Fly-By-Wire flight control system of supersonic jet fighter. The AoA limiter is designed for positive HAoA in longitudinal control law. But, aircraft departure during aggressive negative pitch maneuver such as push over in departure resistance flight test. Therefore negative AoA limiter is needed in longitudinal control law. Result of T-50 flight test show that the AoA is exceed the limit value during aggressive positive pitch maneuver in pull up of power approach mode. In this paper, the AoA limit control law in positive and negative AoA was proposed in order to improve aircraft controllability and stability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.3
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• pp.93-100
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• 2006

Modern versions of supersonic jet fighter aircraft using a digital flight-by-wire flight control system design utilizes a control surface reconfiguration in order to guarantee the aircraft flight stability when a control surface is failed. The T-50 flight control laws are designed such that the surface reconfiguration mode controls the aircraft using non-failed control surfaces when one of the control surfaces is failed. In this paper, linear analysis and HQS(Handling Quality Simulator) pilot simulations are performed to analyze the flight stability and handling quality when the surface reconfiguration mode is engaged for aircraft landing configuration. It is found that the aircraft flight stability and handling quality is satisfied to level 1 requirements when the T-50 flight control law is changed to the surface reconfiguration mode.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.14 no.4
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• pp.1561-1567
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• 2013

This paper is regarding the helicopter flight control law design for the handling quality performance. MIL-F-83300 and ADS-33E specification is used of the helicopter flight handling quality and to meet these requirements, ACAH type controller is required. This paper described the ACAH type controller design and performance evaluations. Helicopter dynamics first developed as nonlinear dynamics including rotor dynamics and then linear model was extracted from hovering to forward flight mode using trim condition. Control law used the model following to meet the handling qualities, the simple inverse model as feed forward gain, decoupling logic and phase model to decouple the axes, and linear model to calculate the coefficients. Handling quality evaluation used the matlab based Conduit tool and verified that Level 1 requirement is satisfied.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.7
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• pp.35-44
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• 2006

When control surface failure occurs, it is conventional to correct a current control or to transform to other control. In this paper, instead of adopting a conventional way, a reconfiguration method which compensate the failure with alternative control surface deflection, depending on the level of failure, by using neural network and PCH(Pseudo-Control Hedging). The Conroller is designed of inner-loop(SCAS : Stability Command Augmentation System) with DMI(Dynamic Model Inversion) and outer-loop with Y axis acceleration feedback for a coordinate turn. Additionally, double PCH method was adopted to prevent actuator saturation and input command was generated to compensate for failure. At the end, The feasibility of the method is validated with randomly selected failure scenarios.

• 제어로봇시스템학회:학술대회논문집
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• 1987.10b
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• pp.537-540
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• 1987

항공기가 평형상태로 비행하는 도중 돌풍과 같은 외부교란을 만난 교란상태운동은 선형화된 미분방정식으로 표현되며 비교적 짧은 비행시간동안의 비행은 선형시 불변계가 된다. 돌풍은 Gauss-Markov확률과정으로 모델링 되며, 항공기가 돌풍을 만난 교란상태운동은 시스템론적으로 보면 백색잡음이 성형필터를 거쳐 계에 입력되는 것과 같다. 초기의 설계방법은 고전적인 주파수영역에서의 해석방법을 사용하였으나 1960년대에 최적제어이론이 도입되면서 평가함수를 사용하여 원하는 비행특성을 얻는 방법을 사용하게 되었다. 그 후 계에 입력되는 외란과 측정시의 잡음으로 인한 불확실한 측정량으로부터 최적상태변수의 추정을 위해 필터링이론을 도입한 확률제어이론을 적용하여 자동조종장치를 설계하게 되었다. 이때까지는 연속제어계로 설계되었으며 그 후 측정신호를 샘플링하여 연속제어계와 등가의 이산제어계를 사용한 자동조종장치가 등장하였으며 이 경우 설계기법으로는 연속제어계를 사용하고 실현시킬 때는 디지털컴퓨터를 사용하였다. 이는 제어하는 동안 계의 계수와 제어법칙을 바꾸어 줄 수 있는 이산제어계의 장점을 이용하지 못하므로 처음부터 계를 등가의 이산계로 보고 제어계를 설계하는 방법이 도입되었다. 이 때 샘플링간격의 결정과 Quantization 영향이 설계시 고려되어야 한다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.11
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• pp.80-87
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• 2005

T-50 supersonic jet trainer aircraft using digital flight-by-wire flight control system receives aircraft flight conditions such as altitude, VCAS(Calibrated Airspeed) and Angle of Attack from IMFP(Integrated Multi-Function Probe). IMFP sensors information have triplex structure using three IMFP sensors. Air-data selection logic is mid-value selection in three information from three IMFP sensors in order to have more reliability. From supersonic flight test at high altitude, air-data information is dropped simultaneously because of supersonic shock wave effect. This error information may affect to aircraft stability and safety in supersonic area at low altitude. This paper propose that sensitivity analysis and HQS(Handling Quality Simulator) pilot simulation in order to analyze flight stability and controllability in supersonic area at low altitude when these information is applied to flight control law.

• Journal of Aerospace System Engineering
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• v.16 no.1
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• pp.17-27
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• 2022

A certain level of stability margin airworthiness criteria should be met to secure robustness against uncertainties between the real plant and the model in a flight control system design. The U.S. Department of Defense (DoD) specification of MIL-F-9490D and airworthiness certification standard of MIL-HDBK-516B uses gain and phase margin criteria of flight control system. However, the same stability margin criteria is applied at all development phases without considering the design maturity of each development phase of the aircraft. Ultimately, a problem arises when the aircraft operation envelope is excessively restricted. This paper proposes the relation of handling qualities and stability margin, and presents re-established stability margin criteria as a development phases and verification methods. The results of the research study are considered to contribute to the verification of the stability margin criteria more flexibly and effectively by applying the method to not only the currently manned developing aircrafts but also the unmanned vehicle to be developed in the future.