The variation of the stress intensity factor (SIF) for cracks growing in integrally machined stiffened skin panels have been studied in an isotropic material. The stringer-skin panel has been modelled using the Nastran finite element system with allowance for crack tip singularity by use of crack tip elements. To allow for variations in stringer and skin geometry, three stiffness ratios (stringer stiffness-skin stiffness) were investigated.

The stress intensity factors have been converted into nondimensionalised values of stress intensity factor coefficients (K1/K0). A damage tolerant and residual strength analysis has been carried out using an in-house computer program. Results indicate very little difference for the different stiffness ratios examined.

A three-dimensional moving mesh method for solving the Euler equations describing the compressible flow about a wing undergoing arbitrary motions and deformations is described. A finite-volume formulation is chosen where the volumes distort as the wing moves or deforms. By using transfinite interpolation, a technique for generating the required sequence of grids has been developed. Furthermore, as the speeds of the grid at the vertices of the finite volumes are required by the flow solver, transfinite interpolation is also used to obtain these by interpolation of the boundary speeds. A two-dimensional version of the method has also been developed and results for both two- and three-dimensional transonic flows are presented and compared with experimental data where available.

At a sufficiently high Reynolds number small separation bubbles, as for example at the leading-edge of an aerofoil, are characterised by laminar flow separation and turbulent reattachment. The transition to turbulence in the separated-flow region through the introduction of a disturbance at the boundary, localised in time and space, is examined. It is found that transition to turbulence is promoted by rapidly growing convective instabilities, and not an absolute instability at the point of application of the disturbance.

This paper describes a concept study of the Spacecab Demonstrator project, which is a small two seat aeroplane capable of routine “zoom-climbs” to space altitude. It is powered by a jet engine and a rocket motor. It would be the first sub-orbital spaceplane since the X-15 of thirty years ago. The study shows that Spacecab Demonstrator should be a simple robust development using existing technology. It would be used as a spaceplane technology testbed, as a reusable sounding rocket for space research and for specialised commercial applications. More importantly, it would serve as a pathfinder for the Spacecab orbital spaceplane, which is designed to be the first fully reusable launch vehicle. Spacecab uses existing engines and proven materials and is designed with provision for certification to airliner standards. Enlarged and mature developments of Spacecab would reduce the cost of sending people into space by three orders of magnitude compared with the cost today, and this could probably be achieved in about 15 years.

The influence of isolated spherical roughness elements upon the transition of a three-dimensional boundary layer on a swept flat plate is discussed in this paper. Perturbations induced by the roughness are found to be more effective on the start of transition rather than the crossflow of the laminar undisturbed boundary layer at the roughness location.

A thin, low cambered aerofoil with a tab at the trailing edge in inviscid, incompressible, two-dimensional, unstalled flow with varying freestream velocity and arbitrary, different motions for both aerofoil and tab is considered. The general expressions for unsteady lift and aerodynamic moment on aerofoil and tab are derived within the assumptions of potential theory.

To verify the approach proposed in this paper, the classical Theodorsen case of an oscillating aerofoil is adopted. For this case, aerofoil loads are calculated in the time domain by applying an inverse Laplace transformation to the approximation of the lift deficiency function in the frequency domain.

The comparison of the results calculated by this method with those obtained by other methods and experiments shows good agreement, which validates the general formulation. The loads on an aerofoil having different frequencies for main aerofoil pitch and tab deflection are calculated.

The results of a blade-vortex interaction (BVI) study incorporating particle image velocimetry (PIV) are presented. Test cases encompass a variety of intersection geometries, including head-on parallel interactions. Results provide quantified flow visualisation images of the flows generated near the blade tip and inboard during interactions. PIV results are examined in the light of understanding gained from previous pressure measurements made from the same BVI facility. Both the difficulties and the advantages of utilising the PIV technique in this application are discussed.

This note presents the results of the application of a momentum control surface analysis to the estimation of the drag of bodies and surfaces placed in the test section of a windtunnel. The assumptions of the method are stated and accuracy of the method considered with particular reference to the drag of riblet surfaces at high subsonic speeds, the drag of two-dimensional cavities at both high subsonic and low supersonic speeds, and the drag of manipulators at high subsonic speeds.

The aerodynamic performance of a multi-element high-lift system has a critical influence on the direct operating cost of a subsonic civil transport aircraft. A thorough understanding of the aerodynamic characteristics of these multi-element aerofoils and wings allows aircraft companies to design and build more competitive aircraft with high-lift systems that are less complex and lighter for given high-lift performance or that have improved lift and drag characteristics for given system complexity and weight. Flight experiments on NASA Langley's B737-100 aircraft have been conducted to further enhance the understanding of the complex flows about multi-element high-lift systems at full-scale flight conditions. In this paper, an overview of the flight program is provided, followed by highlights of experimental results and computational analysis. Measurements included surface pressures on the slats, main element and flap elements using flush pressure ports and pressure belts, surface shear stresses using Preston tubes, off-surface velocity distributions using boundary layer/wake rakes, aeroelastic deformations of the flap elements using an optical positioning system, and boundary layer transition detection using hot-film anemometers and an infrared imaging system. Boundary layer transition measurements on the slat using hot-film sensors are correlated with the flow visualisation results from an infrared imaging technique. Extensive application of several computational techniques and comparisons with flight measurements are shown for a limited number of cases. This program has generated an extensive set of data, much of which are still being analysed.

The flow field of a turbulent free jet issuing from a contoured rectangular nozzle of aspect ratio 2 has been studied experimentally, using hot-wire anemometry. The study was undertaken to gain some understanding of the mixing process within the jet. Results of the measured and derived mean flow and turbulence quantities are presented. The three components of the mean velocity vector, the three Reynolds normal stresses, the two Reynolds primary shear stresses, and the flatness factor of the streamwise fluctuating velocity were measured. Mass entrainment, turbulence kinetic energy, and the intermittency factor were derived from the mean streamwise velocity data, the Reynolds normal stress data, and the flatness factor data, respectively. The derived results and the Reynolds primary shear stress data indicate enhanced near-field mixing of the present rectangular jet compared to a round turbulent free jet. The mean streamwise velocity field changes from a rectangular to an oval and then to an approximately circular shape at about twenty equivalent nozzle diameters downstream of the nozzle exit plane.

Measurements have been made of the surface pressure distributions in the region of the junction between an untapered wing, of NACA 0015 section swept back at 20°, and a flat plate on which a turbulent boundary layer had developed, for several values of wing incidence. It is shown that the section lift coefficient of the wing diminishes, while the pressure drag coefficient increases, as the junction is approached. Oil-flow visualisations on the plate surface show the passage of the horseshoe-like vortex which forms when the retarded boundary layer flow separates as it approaches the leading-edge of the junction. Kinks in the isobars on the plate correlate with the trailing “legs” of this vortex. The surface flow visualisations also show that the turbulence in the junction region spreads onto the wing from the leading edge at an angle of about 10°.

A rotatable X-wire anemometer was used to make measurements of the mean velocity field and of five components of the Reynolds stress tensor, in the wake of the junction, with the wing at incidences of 0° and 9°. Log-law (Clauser) plots were used with the profiles Ū(Y) of longitudinal velocity to estimate the skin friction coefficient on the plate, though adjustments of the zero for Y were necessary to obtain a sensible fit. These corrections were often larger than can be readily explained, but the skin-friction values are consistent with the corresponding, measured velocity correlation, . The Reynolds stresses in the wake region clearly show that the horseshoe-like vortex legs persist beyond the trailing-edge of the wing, the turbulence intensity being larger on the suction side for the wing at incidence.

Average values of the skin-friction coefficient on the plate in the junction region are little different from those away from the junction. No corresponding information is available for the skin-friction on the wing, but if this is likewise unaffected by the junction, then the total drag of the junction region will be greater than the sum of those of the isolated parts, simply because of the increase in pressure drag.

Effective control of combat aircraft at high angles of attack requires the flight control system to operate to higher frequencies and gains. This increases the possibility of excitation of the structural modes by altering the aeroservoelastic stability of the complete system. This paper details the results of tests carried out on a rigid, 55° leading edge sweep delta wing, half model fitted with a half span elevon operated through a closed loop control system. The effects of steady elevon deflections up to 40° angle of attack are described. The unsteady pressures resulting from elevon oscillations at frequency parameters up to a value equivalent to the first wing bending mode of a typical combat aircraft have been measured. The effect of the elevon oscillation on the elevon itself and on the bending mode have been estimated by calculating the unsteady hinge and root bending moments about the wing root. The lower surface, unsteady elevon effects were found to be independent of incidence and frequency parameter. The amplitude of response of the weak vortex on the upper surface of the wing reduced rapidly with frequency parameter. Above an incidence of 10°, the vortex breakdown was above the wing and led to a further reduction in the vortex response. The vortex response lagged the elevon motion.

An experimental investigation has been made into the effects of riblets on the drag in a rectangular duct of aspect ratio 10. It has been found that a maximum drag reduction of approximately 10% occurs in fully developed turbulent flow. This figure is larger than the 6–8% drag reduction usually found for external flow. The maximum drag reduction occurred when the height of the riblets was between 7 and 13 wall units. Analysis of the developing region of the flow has shown a maximum drag reduction of approximately 8% at a wall unit value of 13, consistent with results in external flows. The effects of riblets on transition has also been investigated. It was found that riblets delayed transition by 2–4% in critical Re as measured by peak flatness, and caused some extension in the length of the transition process, so that the completion of transition was delayed by between 4–8%, with the stronger effect at the lower Reynolds numbers.

This paper describes an investigation of the aerodynamic characteristics of collectively variable incidence wing tips at low speeds. Theoretical considerations supported by experimental measurements suggest that the use of articulated tips which can be rotated to a high incidence for low speed flight will decrease sensitivity to gusts. Furthermore, windtunnel measurements demonstrate that the use of very high tip incidences can produce an increment in maximum overall lift coefficient for some wing configurations. Another feature investigated is the possibility of utilising the lift increment for the purposes of direct lift control. The use of a radio controlled model to test the practical viability of the concept is described.

The equivalent continuum beam-rod model of an aircraft wing structure with composite laminated skins has been developed based on the concept of energy equivalence. The equivalent structural properties of the continuum beam-rod model are obtained by directly comparing the reduced stiffness and mass matrices for a typical segment of aircraft wing with those for a finite element of continuum beam-rod model. The finite element stiffness and mass matrices are condensed through the well known finite element formulation procedure to be used in calculating the reduced stiffness and mass matrices for the aircraft wing segment in terms of the continuum degrees of freedom introduced in this paper. The present method of continuum modelling may yield every equivalent structural property including all possible couplings between bending, torsional and transverse shear deformations. To evaluate the equivalent continuum beam-rod model developed herein, the free vibration and aeroelastic analyses for a box-beam type aircraft wing structure have been conducted by using the continuum beam-rod model, and the numerical results are compared with the results by using other different models of the aircraft wing structure.

Prior to the adoption of modern flybywire high authority flight control systems, the usable lift or angle of attack was usually dictated by the wing stall and subsequent loss of control of the aircraft.

With the advent of the high authority FCS, with high levels of integrity, the designer of modern combat aircraft faces the possibility of providing good levels of aircraft control up to and well beyond the AoA at which maximum lift or loss of control would have occurred in the past. Further, the use of such control systems has allowed maximum advantage to be taken of control configured vehicle technology to enhance the performance of the aircraft within the conventional angle of attack range.

This paper provides a review of the experience which has been gathered within BAe Military Aircraft, Warton, in the use and benefits of these technologies with regard to the high AoA capability and impact on maximum usable lift. Examples are drawn from the Tornado, Jaguar FBW and EAP aircraft programmes, all of which have employed flybywire control, to some degree or other, combined with varying degrees of CCV technology, leading to their adoption on the Eurofighter 2000 programme.

In doing this, some of the engineering compromises which the modern aircraft designer faces in achieving the customer required levels of aircraft performance will be addressed, together with the outcome which has satisfied these purposes in the particular example considered. In particular, the power of the modern FCS places an increased reliance on the knowledge of the aerodynamics of the aircraft, around which the FCS must be designed. This raises particular challenges in the increased AoA ranges which associate with the maximum lift of these high performance configurations.

The paper concludes with an indication of the developments in technology which are evolving and which will lead to further advances and increased capability combat aircraft in future years.

This note reports tests in a shock tunnel in which a fully integrated scramjet configuration produced net thrust. The experiments not only showed that impulse facilities can be used for assessing thrust performance, but also were a demonstration of the application of a new technique(1) to the measurement of thrust on scramjet configurations in shock tunnels. These two developments are of significance because scramjets are expected to operate at speeds well in excess of 2 km/s, and shock tunnels offer a means of generating high Mach number flows at such speeds.

The dynamic stability derivatives of blunt cones for small variations in angles of attack have been previously derived by the current author. However, no account of the unsteady nature of the boundary layer was made in that work. In this paper closed form expressions for the increment in dynamic stability due to the presence of the boundary layer are derived by considering the pressure distribution perturbations as the boundary layer continuously adjusts to the enclosed oscillating body. The theory provides a first hand estimate of the complete pitch derivative damping without having to resort to more rigorous and expensive computational methods. Calculations performed at Mach numbers of 7 to 10 with axis positions ranging from 0·5 to 0·7 of the chord length for cones of semi-angle 10° and 20°, indicate that the effect of boundary layer is to slightly reduce the magnitude of the inviscid damping derivative. For blunt cones at angles of attack less than 5°, this was in good agreement with the limited experimental data available.

The results of a computational fluid dynamics analysis of three prominent World War II fighter aircraft, the Supermarine Spitfire, North American P-51 Mustang and Focke Wulf Fwl90 are presented. The goal of this analysis was to learn the strengths and weaknesses inherent in the aerodynamics of these aircraft. Design features covered include comparisons of the wing designs and calculated total wetted areas along with a look at specific details of the aerodynamics of each aircraft. This information, while historical, still has relevance in today's world of aircraft design.

This paper concerns the study of the three-dimensional interaction of a slat wake and a wing boundary layer, as commonly occurs on the suction surface of a high-lift rnulti-element aerofoil. The experimental test case has a realistic slat-wing flow configuration under infinite swept wing conditions. This enables the effect of the cross flow induced by the wing sweep on the slat wake and the wing boundary layer interaction process to be investigated. A triple hot-wire system has been used to measure the mean velocity and Reynolds stress components in the flow field. The measurements are used to validate a three-dimensional boundary layer code, which has been modified to calculate this three-dimensional wake-boundary layer mixing flow.