The class of aircraft known as gyroplanes (or autogyros) helped to pave the way for the development of the helicopter. However, they have found no application in contemporary commercial or military aviation. It is in recreational or sport flying that the gyroplane has proved popular. Most if not all designs are however homebuilts, and as a consequence little analysis of any significance has been conducted on the flight mechanics of these aircraft. This Paper presents an analysis of the longitudinal stability of gyroplanes. Simple consideration of the basic forces and moments that might influence stability, in the context of elementary rotor behaviour, is used to assess the fundamental nature of gyroplane longitudinal stability. This is quantified by the application of a sophisticated generic rotorcraft mathematical model to the gyroplane problem. It is concluded that the basic configuration can be longitudinally stable with appropriate design, and that the rotorspeed degree of freedom must also be included when modelling the aircraft.

“Urban air vehicles” have been hailed as the next revolution in aviation. Prototypes of various sizes have been flown to demonstrate basic flight (hover and climb), but in most cases there is no demonstration of full flight capability, for example conversion from vertical to level flight (conversion corridor). There are proposals for vehicles in a wide range of scales: from drones specifically designed to deliver goods, to full size vehicles for manned transportation. Most of the concepts proposed include full electric propulsion, multiple (often convertible) rotors (ducted or un-ducted, counter-rotating), and widespread use of composite materials. Start-up companies are seeking funding with high-profile demonstrations in front of the media, but many unresolved technical problems are not been solved. Large aerospace companies have joined the fray. These initiatives are fuelling expectations that achieving the next milestone is within easy reach. This paper aims to fill some gaps in understanding and curb optimism. It takes a holistic view in order to establish a scientific basis for design, manufacturing, operations.

An analysis technique capable of simulating the effect of engine failure during takeoff from offshore platforms is presented. Use is made of inverse simulation whereby the control actions required for a subject helicopter to follow a particular trajectory can be established. A mathematical representation of the towering takeoff procedure, and details of the modified inverse simulation technique needed to cope with the modelling of an engine failure are described. Detailed piloting accounts of the strategy used to fly the towering takeoff (with and without engine failures) are also given and are used to give qualitative validation of the analytical approach. Simulation results for a single engine failure of a transport helicopter during critical phases of a towering takeoff are presented. Finally, some directions for future work and potential applications of the technique are discussed.

A class of delta wings is considered, whose under-surface has an inverted-V, or inverted-W, cross section of such a form that, at the “design” Mach number and incidence, the shock waves formed are plane. The geometry of the shock-wave system and surface is described briefly, and comments made about the utility of the concept in relation to hypersonic flow studies.

The solution of a set of m ordinary linear differential equations of the nth order having variable coefficients can be provided by the use of a set of m2 indicial admittances and the solution is derived from that given by the use of impulsive admittances. The indicial admittances are easily derived from the impulsive admittances and vice versa.

An “indicial admittance” is a function representing the response of a system to an applied “input” in the form of a “unit step function” and the concept is due to Heaviside. When the indicial admittance is known, the response to an input which varies with time in an arbitrary manner can easily be found. Similarly, the response can easily be found when the “impulsive admittance” is known, where this is the response to a unit instantaneous impulsive input.

Results of a comparison of dynamic stall onset are presented, as assessed from low Mach number windtunnel data contained in the University of Glasgow's database and that from the well established Beddoes’ model. The model, which was originally developed to reconstruct higher Mach number dynamic stall characteristics, exhibited a lower stall onset incidence than that assessed from the windtunnel data. The differences and speculations on the physical reasoning underlying the two assessments are discussed. An addition to the Beddoes model is proposed which yields an improved reconstruction of the Glasgow windtunnel data.

A conventional inverse simulation does not accommodate control constraints; hence for aggressive manoeuvring flight conditions, where control inputs are close to the limits, these algorithms lose some of their applicability. A modification of the conventional inverse simulation technique that accommodates the onset of physical limits or constraints is proposed in this paper. In this way a process of constraints handling is incorporated into the inverse simulation algorithm. Therefore, the aim of this paper is to demonstrate that conventional inverse simulation can be improved in terms of the realism of the results by applying a predictive capability for applications involving manoeuvring flight. The paper gives details of the development of the predictive inverse simulation algorithm and helicopter model used and, by presenting examples of results calculated for pop-up and lateral realignment manoeuvres demonstrates that a ‘receding horizon’ predictive approach offers improvements in the realism of inverse simulation results.

The coaxial compound helicopter with lift-offset rotors has been proposed as a concept for future high-performance rotorcraft. This helicopter usually utilizes a variable-speed rotor system to improve the high-speed performance and the cruise efficiency. A flight dynamics model of this helicopter associated with rotor speed governor/engine model is used in this article to investigate the effect of the rotor speed change and to study the variable rotor speed strategy. Firstly, the power-required results at various rotor rotational speeds are calculated. This comparison indicates that choice of rotor speed can reduce the power consumption, and the rotor speed has to be reduced in high-speed flight due to the compressibility effects at the blade tip region. Furthermore, the rotor speed strategy in trim is obtained by optimizing the power required. It is demonstrated that the variable rotor speed successfully improves the performance across the flight range, but especially in the mid-speed range, where the rotor speed strategy can reduce the overall power consumption by around 15%. To investigate the impact of the rotor speed strategy on the flight dynamics properties, the trim characteristics, the bandwidth and phase delay, and eigenvalues are investigated. It is shown that the reduction of the rotor speed alters the flight dynamics characteristics as it affects the stability, damping, and control power provided by the coaxial rotor. However, the handling qualities requirements are still satisfied with different rotor speed strategies. Finally, a rotor speed strategy associated with the collective pitch is designed for maneuvering flight considering the normal load factor. Inverse simulation is used to investigate this strategy on maneuverability in the Push-up & Pull-over Mission-Task-Element (MTE). It is shown that the helicopter can achieve Level 1 ratings with this rotor speed strategy. In addition, the rotor speed strategy could further reduce the power consumption and pilot workload during the maneuver.

Significant progress has been made to date in modelling, computationally, the formation and development of the dust cloud that forms in the air surrounding the rotorcraft under brownout conditions. Modern computational methods are able to replicate not only the development of the dust cloud in appropriate operational scenarios, but also the sensitivity of the shape and density of the dust cloud to the detailed design of the rotorcraft. Results so far suggest that attempts to ameliorate brownout by aerodynamic means, for instance by modifying the rotor properties, will be frustrated to some extent by the inherent instability of the flow field that is produced by the helicopter. Nonetheless, very recent advances in understanding the fundamental mechanisms that lead to the formation of the dust cloud may allow substantial progress to be made once certain elements of the basic physics of the problem are more fully understood and better quantified.

A method is given for finding the reciprocal of a matrix which is triply partitioned horizontally and vertically in such a manner that the sub-matrices in the principal diagonal are square, but these matrices need not be of the same order. A preliminary rearrangement of the matrix may be helpful.

In theoretical work it is sometimes required to find the reciprocal of a matrix which can be so partitioned that some of the sub-matrices are of simple types, e.g. triangular or nul matrices, and the calculation of the reciprocal may then be facilitated by using formulae for the inversion of partitioned matrices. The same method will be often advantageous in numerical work also. In an earlier paper the reciprocation of doubly-partitioned matrices was treated in a general way and a method was given for triply-partitioned matrices subject to the restriction that the sub-matrices in the principal diagonal are square and of the same order.

This paper presents a method for assessing two-dimensional aerofoil lift and pitching moment characteristics including trailing edge and gross laminar separation. The model used is a direct viscid-inviscid interaction scheme based on a vortex panel method with boundary-layer corrections and an inviscidly modelled wake. The integral boundary-layer methods adopted behave well in the region of separation and thus, good comparisons with measured separation characteristics are obtained. Generally the predictions of lift and pitching moment may be considered to be within the experimental error, but where this is not the case, the applicability of the modelling technique is discussed.

This paper examines the dynamic stalling of three wing planforms and characterises the main features of the stalling process in each case. The particular data were obtained during a three year research programme in the Department of Aerospace Engineering, University of Glasgow to collect high-resolution unsteady pressure data on the dynamic stalling characteristics of finite wing planforms. In this study, which was motivated by the pressing need for a greater understanding of the strongly three-dimensional effects in the tip region of helicopter rotors, the three wing planforms considered were a straight rectangular wing, a rectangular wing with swept tips and a delta wing. The initial test programme was followed by a further three years of detailed analysis and interpretation of the test data. Results from this analysis are presented in the present paper for cases in which the wings were subject to ramp motions.

A symmetric Mach 12·76 hypersonic flow in a 90° corner, formed by 30° swept back intersecting 8° wedges, was investigated in detail through numerical simulation using locally conical Navier-Stokes equations. Three different numerical schemes for spatial discretisation, MacCormack central differencing, van Leer’s flux vector splitting and Osher’s flux difference splitting, were studied to compare their capabilities to capture both strong shock waves and thin shear layers. Comparison with experimental data was made to validate the simulation. The numerical simulation provided further insight into the flowfield and a pair of counter-rotating vortices were discovered near the junction of the corner.

Numerical analysis of the flow in weapon bays modelled as open rectangular cavities of length-to-depth (L/D) ratio of 5 and width-to-depth (W/D) ratio of 1 with doors-on and doors-off is presented. Flow conditions correspond to Mach and Reynolds numbers (based on cavity length) of 0·85 and 6·783m respectively. Results from unsteady Reynolds-averaged Navier-Stokes (URANS), large-eddy simulation (LES) and detached-eddy simulation (DES) are compared with the simulation methods demonstrating the best prediction of this complex flow. It was found that URANS was not able to predict the change of flow characteristics between the doors-on and doors-off configurations. In addition, the energy content of the cavity flow modes was much better resolved with DES and LES. Further, the DES was found to be quite capable for this problem giving accurate results (within 3dB of) experiments and appears to be a promising alternative to LES for modelling massively separated flows.

Unsteady spiked body flows were simulated by a second order time-accurate CFD method. Laminar, axisymmetric flow was considered at Mach 2.21 and Mach 6 freestreams and Reynolds’ numbers based on the blunt body diameter of 0.12 million and 0.13 million, respectively. A spiked forward facing cylinder with spike lengths between LID = 1.00 and LID = 2.40 was used as the model geometry. Following the numerical method’s verification, the individual flow modes of oscillation and pulsation were examined. The frequency of the events was found in good agreement with the experiment, while the pressure amplitudes were overpredicted in the Mach 6 cases. Analysis of the numerical results showed that the oscillation flow mode was driven by a viscous mechanism, whereas the pulsation by an inviscid one. The hysteresis phenomenon in the transition between the two flow modes was predicted qualitatively.

The experimental results of an oblique Blade-Vortex Interaction (BVI) study are presented. The quality of all pressure data reflects improvements in the Glasgow University BVI facility and in the method of reducing and presenting data. The data collected during oblique interactions is found qualitatively and quantitatively similar to that collected in corresponding parallel interactions, for interactions within ± 30° of parallel. Details of the pressure data are examined in the light of understanding gained from parallel BVI experimentation. The study highlights the effects of three dimensional flow interactions not accounted for in the nondimensionalisation traditionally utilised in the analysis of parallel blade-vortex interactions.

A recent study by digital computer of the theoretical characteristics of a family of low drag aerofoil sections was prompted some time ago by the need to redesign a particular NACA 6-series aerofoil of high camber (which was not producing the desired pressure distribution). Understandably, therefore, this new family of aerofoils—christened the GU series—was patterned on the NACA sections, and designed to provide the characteristic region of uniform velocity over the forward portion of one surface at the extreme of the low drag range, followed by a compression region of roughly constant adverse pressure gradient.

This paper evaluates a time marching simulation method for flutter which is based on a solution of the Euler equations and a linear modal structural model. Jameson’s pseudo time method is used for the time stepping, allowing sequencing errors to be avoided without incurring additional computational cost. Transfinite interpolation of displacements is used for grid regeneration and a constant volume transformation for inter-grid interpolation. The flow pseudo steady state is calculated using an unfactored implicit method which features a Krylov subspace solution of an approximately linearised system. The spatial discretisation is made using Osher’s approximate Riemann solver with MUSCL interpolation. The method is evaluated against available results for the AGARD 445.6 wing. This wing, which is made of laminated mahogany, was tested at NASA Langley in the 1960s and has been the standard test case for simulation methods ever since. The structural model in the current work was built in NASTRAN using homogeneous plate elements. The comparisons show good agreement for the prediction of flutter boundaries. The solution method allows larger time steps to be taken than other methods.

Potential function methods have been applied extensively to many terrestrial and space robot control problems. This paper extends these methods to multiple robot problems for on-orbit assembly. By considering a population of free-flying robotic vehicles in orbit, capable of both rotation and translation, the potential function method reduces complex assembly problems to sets of simple translation and rotational commands. This is achieved using a potential function incorporating both collision avoidance between the free-flyers and connection constraints in the assembled structure. The method also allows a subsumptive type control architecture with the flexibility to carry out many different assembly tasks simultaneously.

As propeller-driven aircraft are the best choice for short/middle-haul flights but their acoustic emissions may require improvements to comply with future noise certification standards, this work aims to numerically evaluate the acoustics of different modern propeller designs. Overall sound pressure level and noise spectra of various blade geometries and hub configurations are compared on a surface representing the exterior fuselage of a typical large turboprop aircraft. Interior cabin noise is also evaluated using the transfer function of a Fokker 50 aircraft. A blade design operating at lower RPM and with the span-wise loading moved inboard is shown to be significantly quieter without severe performance penalties. The employed Computational Fluid Dynamics (CFD) method is able to reproduce the tonal content of all blades and its dependence on hub and blade design features.