This paper describes the development and validation of a high fidelity simulation model of the Bell 412 helicopter for handling qualities and flight control investigations. The base-line model features a rigid, articulated blade-element formulation of the main rotor, with flap and lag degrees of freedom. The Bell 412 HP engine/governor dynamics are represented by a second-order system. Other key features of the base-line model include a finite-state dynamic inflow model and lag damper dynamics. The base-line model gives excellent agreement with flight-test data over the speed range 15-120kt for on-axis responses. Prediction of off-axis responses is less accurate. Several model enhancement options were introduced to obtain an improved off-axis response. It is shown that the pitch/roll off-axis responses in transient manoeuvres can be improved significantly by including wake geometry distortion effects in the Peters-He finite-state dynamic inflow model.

Our paper concerns the severity of the response of a rotorcraft encountering the vortex of a fixed-wing aircraft. One of the key questions is whether a rotorcraft designed to meet handling performance standards will have sufficient control margin for the pilot to overcome the effects of a vortex encounter. This question is addressed through an analytical study supported by preliminary piloted simulation tests. The handling criteria are found to be well suited to establishing the severity and associated hazard category of typical encounters. Cases are illustrated where insufficient control margin is available to overcome the effects of the encounter. The pilot intervention time is critical as expected. In risk assessment parlance, an intervention time of 3 seconds leads to hazardous encounters (> Level 3 HQs) while a reduced intervention time of 1.5 seconds is more likely to be major (Level 3 HQs). List of Symbols p, q, r roll, pitch and yaw rates qpk peak pitch rate Q = qpk /∆θ pitch attitude quickness r vortex radial dimension rc vortex core radius VT(r) vortex tangential velocity Vc vortex core velocity θ, φ, ψ pitch, roll and yaw angles Γ ortex circulation 25-1

NOMENCLATURE k variable describing profile of motion [-] p, q, r angular roll, pitch and yaw rates in aircraft body axes [deg/s] Q e engine torque [Nm] t time [s] T total duration time of motion [s] x, y, z position [m] Greek notation φ, θ, Ψ Euler angles, describing aircraft attitude [deg] θ 0 , θ 0,tr main rotor and tail rotor collective pitch [deg] θ 1s , θ 1s lateral and longitudinal cyclic pitch [deg] χ gap [m, deg, N, etc.] τ time to close a gap [s] Ω main rotor rotational speed [%] Χ state [m, deg, etc.] ABSTRACTResearch studies have indicated that the optical flow parameter, time to close tau, is the basis of purposeful control in the animal world, and used by both fixed wing and helicopter pilots during manoeuvring. This parameter is defined as the instantaneous time to close a gap (spatial or force) at the current closing rate. A novel automatic flight control strategy has been developed that makes use of optical flow theory and in particular, the parameter tau. This strategy has been applied to two distinct problems; (1) the landing of a helicopter on a ship and (2) the lateral repositioning of a helicopter. The first is a challenging case because the landing of a helicopter on a ship is one of the most dangerous of all helicopter flight operations. Furthermore, helicopters are often subject to torque oscillations during rapid collective control, which increases pilot workload significantly when operating with low power margins and/or whilst performing tasks that require accurate heave control. The second case demonstrates the generality of the technique. Both automatic manoeuvres were simulated successfully within desired limits, with the novel control strategy creating a 'natural', smooth, tau motion.

This paper presents the first results from research into active control of structural load alleviation (SLA) for tiltrotor aircraft carried out in the European 'critical technology' RHILP project. The importance of and the need for SLA in tiltrotors are discussed, drawing on previous US experience reported in the open literature. The paper addresses the modelling aspects in some detail; hence forming the foundation for both the FLIGHTLAB simulated XV-15 and EUROTILT configurations. The primary focus of attention is the suppression of in-plane rotor yoke loads for pitch manoeuvres in airplane mode; without suppression these loads would result in a very high level of fatigue damage. Multi-variable control law design methods are used to develop controller schemes and load suppression of 80-90% is demonstrated using rotor cyclic control, albeit at a 20-30% performance penalty. However, rotor flapping transients tend to increase by the action of the SLA system. A dual-objective control design approach demonstrates the effectiveness of suppressing both loads and flapping simultaneously.

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