The paper focuses on the modelling, simulation and control of an electrical miniature actuator integrated in the actuation mechanism of a new morphing wing application. The morphed wing is a portion of an existing regional aircraft wing, its interior consisting of spars, stringers, and ribs, and having a structural rigidity similar to the rigidity of a real aircraft. The upper surface of the wing is a flexible skin, made of composite materials, and optimised in order to fulfill the morphing wing project requirements. In addition, a controllable rigid aileron is attached on the wing. The established architecture of the actuation mechanism uses four similar miniature actuators fixed inside the wing and actuating directly the flexible upper surface of the wing. The actuator was designed in-house, as there is no actuator on the market that could fit directly inside our morphing wing model. It consists of a brushless direct current (BLDC) motor with a gearbox and a screw for pushing and pulling the flexible upper surface of the wing. The electrical motor and the screw are coupled through a gearing system. Before proceeding with the modelling, the actuator is tested experimentally (stand alone configuration) to ensure that the entire range of the requirements (rated or nominal torque, nominal current, nominal speed, static force, size) would be fulfilled. In order to validate the theoretical, simulation and standalone configuration experimental studies, a bench testing and a wind-tunnel testing of four similar actuators integrated on the real morphing wing model are performed.

Box-like slung loads exhibit periodic yaw response instabilities, while carried externally by a helicopter. When coupled with the slung load longitudinal and lateral pendulum motions, these instabilities result in significant pendulum oscillations of the load. High amplitude oscillations lead in many cases to the limiting of a load’s flight envelope. Using wind tunnel and flight tests, rear mounted fins were previously demonstrated as efficient means for stabilisation of a problematic load. However, the lack of a proper analytical model of the stabilised load’s aerodynamic characteristics, led to a trial and error development process, without an appropriate physical understanding of the stabilisation problem. The present paper describes a method for the aerodynamic modeling of fins stabilised slung loads based on a limited number of simple static wind-tunnel tests. The resulting database is incorporated in a dynamical slung load simulation that shows good agreement with dynamic wind-tunnel tests. The applicability of the proposed method is demonstrated, by the calculation of stabilised loads aerodynamic databases for interim fin inclination angles not covered by tests.

The terminology ‘rotorcraft-pilot coupling’ denotes phenomena arising from interaction between pilot and rotorcraft. Among these, the present work deals with ‘pilot-assisted oscillations’ that derive from unintentional pilot actions on controls due to seat vibrations, and are strictly related to rotor-aeroelasticity/airframe-structural-dynamics coupling, with involvement of blade control actuator dynamics. Focusing the attention on helicopters, a comprehensive rotorcraft model is developed and applied, with main rotor unsteady aerodynamics described in state-space form. This makes it particularly suited for stability and frequency-response analysis, as well as control applications. Numerical investigations address two critical rotorcraft-pilot coupling aeroelastic issues: stability of vertical bouncing and gust response in hovering. Results from main rotor unsteady aerodynamics modelling are compared with widely-used quasi-steady aerodynamics predictions. These suggest that, for accurate RPC/PAO phenomena predictions, mathematical modelling should include the three-dimensional, unsteady-flow effects, and that the pilot-in-the-loop passive behaviour produces a beneficial effect on the load factor generated by gust encountering.

When an aero-engine is operating on the ground the formation of a potent inlet ground vortex can arise which has the ability to ingest foreign object debris. The ingestion of foreign objects can cause notable damage to engine components as well as overall performance degradation. The assessment of foreign object ingestion has been conducted using a combination of computational fluid dynamics, analytical modelling and an Euler-Lagrange uncoupled discrete phase particle tracking method. The flow fields for a full-scale aero-engine have been simulated for a range of ground clearances and intake velocity ratios under crosswind conditions. The sensitivity of the debris ingestion thresholds and characteristics to particle size and material density has been evaluated. The characteristics of the ingestion location within the intake are also considered for a range of operating conditions and particle parameters.

Military aircrafts are often subjected to severe flight maneuvers with high Angles-of -Attack (AOA) and Angles of Sideslip (AOSS). These flight attitudes induce non-uniform in flow conditions to their gas turbine engines which may include distortion of inlet total pressure and total temperature at the Aerodynamic Interface Plane (AIP). Operation of the downstream compression system with distorted inflow typically results in reduced aerodynamic performance, reduced stall margin, and increased blade stress levels. In the present study the steady state total pressure distortion induced to the Aerodynamic Interface Plane due to the aircraft’s flight attitude have been estimated in terms of distortion descriptors. The distorted conditions at the interface between the intake and the engine have been predicted by using Computational Fluid Dynamics (CFD), where 33 different aircraft flight attitudes have been tested. Based on the obtained results the effect of Angle-of-Attack (AOA) and Angle of Side Slip (AOSS) on the distortion descriptors have been studied. The results showed that the distortion effect becomes more pronounced whenever this specific airframe configuration is exposed to incoming flow with an AOSS. Among the tested cases, the greatest total pressure defect at the AIP in terms of difference from the average value and of circumferential extent was calculated for the flight attitudes of 0·35M flight with 0° AOA and 8° AOSS and 0·35M fight with 16° AOA and 16° AOSS.

Estimation of the lift of an aerofoil is one of the fundamental measurements of fluid mechanics. Lift is commonly measured using a load cell or a force balance. Non-intrusive methods to measure lift are usually pressure based. Aerofoils may be pressure tapped where small surface orifices are connected via tubing to a pressure measurement system, either a multi-tube manometre or an electronic system. Both measurement options add cost and complication, especially in an educational setting. Pressure tapping small aerofoils can also be difficult, especially if the models are rapid prototyped (RP). Low model surface resolution (from RP manufacture) and confined geometry complicate model assembly and finishing. Boundary-layer transition caused by poorly implemented tappings (too large a diametre or poorly aligned, i.e. straight aft) can also alter results. Wall pressure tappings may also be used and have the benefit of being non-intrusive. To implement, the test section roof and floor is tapped with a streamwise row of ports that facilitate measurement of the wall pressure signature. Integration of the pressure differential then relates to the lift produced. This measurement methodology still requires a multi-channel pressure acquisition system and modification of the wind tunnel. In Refs 4,5 methods are presented that facilitate calculation of the instantaneous forces acting on a body through flow field measurements determined using particle image velocimetry. However, the required flow field measurements encompass those surrounding the body, and are not a simple point measurement. In Ref. 6 a method is presented to estimate the lift of an aerofoil using two Pitot-static tubes that are used to measure the velocity above and below the aerofoil’s quarter chord. Wall corrections are required to yield an accurate lift estimate.