This paper presents results from a joint Lockheed Martin/NASA Glenn effort to design and verify an ultra-compact, highly-survivable engine inlet subsonic duct based on the emerging technology of active inlet flow control (AIFC). In the AIFC concept, micro-scale actuation (∼mm in size) is used in an approach denoted ‘secondary flow control’ to intelligently alter a serpentine duct's inherent secondary flow characteristics with the goal of simultaneously improving the critical system-level performance metrics of total pressure recovery, spatial distortion, and RMS turbulence. In this approach, separation control is a secondary benefit, not a design requirement. The baseline concept for this study was a 4:1 aspect ratio ultra-compact (LID= 2·5) serpentine duct that fully obscured line-of-sight view of the engine face. At relevant flow conditions, this type of duct exhibits excessive pressure loss and distortion because of extreme wall curvature. Two sets of flow control effectors were designed with the intent of establishing high performance levels to the baseline duct. The first set used two arrays of 36 co-rotating microvane vortex generators (VGs); the second set used two arrays of 36 micro air-jet (microjet) VGs, which were designed to produce the same ‘vorticity signature’ as the microvanes. Optimisation of the microvane array was accomplished using a design of experiments (DOE) methodology to guide selection of parameters used in multiple Computational Fluid Dynamics (CFD) flow solutions. A verification test conducted in the NASA Glenn W1B test facility indicated low pressure recovery and high distortion for the baseline duct without flow control. With microvane flow control, at a throat Mach number of 0·60, pressure recovery was increased 5%, and both spatial distortion and turbulence were decreased approximately 50%. Microjet effectors also provided significantly improved performance over the baseline configuration.

Direct Numerical Simulation (DNS) results are presented for high speed nonreacting mixing layer in a confined test section. The hyper-velocity mixing layer experiment of Erdos et al with H2/N2 stream is simulated by discretizing two dimensional Navier Stokes equation using a higher order (fourth order spatial and second order temporal) compact numerical algorithm. A favourable comparison of the computation with experimentally measured wall static pressure forms the basis of further analysis. Instantaneous flow picture and the mean profiles of various flow variables were examined to determine the development and general characteristics of the confined mixing layer. It has been found that the growth of the mixing layer is towards the high speed side of the layer. Various turbulence quantities were derived from the stored time series data of the DNS calculation and the results were compared with the experimental results of supersonic free shear layer as no experimental results of turbulence statistics are available for the confined hypervelocity mixing layer. The increasing Reynolds stress data with the flow direction indicate that the turbulence is sustained by transferring the energy from the mean flow to the fluctuating field as the shear layer develops. Although the Reynolds stress is negligible in the most portion of the wall boundary layers, effect of counter gradient effect is observed in the far downstream location of the lower wall boundary layer. The general conclusion that for the supersonic mixing layer, various turbulence quantities like Reynolds stress, turbulence intensities (both streamwise and transverse) decrease with the increase in the convective Mach number is also confirmed by our results.

Investments in aeronautics research and technology have declined substantially over the last decade, in part due to the perception that technologies required in aircraft design are fairly mature and readily available. This perception is being driven by the fact that aircraft configurations, particularly the transport aircraft, have evolved only incrementally over the last several decades. If, however, one considers that the growth in air travel is expected to triple in the next 20 years, it becomes quickly obvious that the evolutionary development of technologies is not going to meet the increased demands for safety, environmental compatibility, capacity, and economic viability. Instead, breakthrough technologies will be required both in traditional disciplines of aerodynamics, propulsion, structures, materials, controls and avionics as well as in the multidisciplinary integration of these technologies into the design of future aerospace vehicles concepts. The paper discusses challenges and opportunities in the field of aerodynamics over the next decade. Future technology advancements in aerodynamics will hinge on our ability to understand, model and control complex, three-dimensional, unsteady viscous flow across the speed range. This understanding is critical for developing innovative flow and noise control technologies and advanced design tools that will revolutionise future aerospace vehicle systems and concepts. Specifically, the paper focuses on advanced vehicle concepts, flow and noise control technologies, and advanced design and analysis tools.

Using an exact solution of the Euler equations, the pressure distributions over some axisymmetric bodies with a negative pressure gradient over the majority of the surface (from 99.7% to 88% of the total body length) have been calculated. The results of wind tunnel tests for these bodies are presented.

This paper considers defence and aerospace research in the UK from the perspective of a research worker. The distinction between research, technology demonstration and development, the particular characteristics of research and the qualities required in a research worker are discussed. The important influence of the research environment, both within an organisation and nationally, is highlighted. The contribution of the past environment to the research underlying the present strength of UK technology, and subsequent adverse changes in the environment, are noted. The decline of UK research funding and of the UK research base is contrasted with the situation of our main industrial competitors. Finally, the author lists the actions which he believes are needed to rebuild research to a level which will underwrite the long-term future of UK defence and aerospace.

This paper investigates the potential benefits of utilising bifurcation analysis results in the planning and evaluation of piloted simulator trials of manoeuvrable aircraft, particularly when encountering nonlinear phenomena such as departure. Comparisons of bifurcation diagram predictions with piloted simulations are performed in order to validate the technique for application to realistic flight conditions. A close relationship is found between the theoretical and the experimental responses, and potential improvements in effectiveness of piloted test programmes are identified. Proposals for enhancing the predictive capability of bifurcation methods are made.

Fibre reinforced metal laminates (FRMLs) are a new family of aerospace structural materials developed for fatigue critical applications. These materials are laminated sheets of thin and high strength metal layers, usually aluminium, and alternating plies of fibre reinforced polymer composite layers, viz, aramid/epoxy, carbon/epoxy, glass /epoxy, etc. Afaghi-Khatabi and Ye have considered the fracture data existing on various ARALL (aramid reinforced aluminium) laminates containing a circular hole for evaluation of notched tensile strength, using the effective crack growth model (ECGM), point stress criterion (PSC) of Whitney and Nuismer and damage zone criterion (DZC) of Eriksson and Aronsson. It is noted that ECGM overestimates, whereas PSC and DZC underestimate the notched tensile strength of FRMLs. This note highlights briefly the two different criteria of Whitney and Nuismer known as ‘point stress criterion (PSC)’ and ‘average stress criterion (ASC)’, and suggests a modification in the criteria for accurate evaluation of the notched tensile strength.

'Better, faster, cheaper’ (BFC) emerged in the 1990s as a new paradigm for aerospace products. In this paper, we examine some of the underlying reasons for BFC and offer some thoughts to help frame the thinking and action of aerospace industry professionals in this new era. Examination of literature on industrial innovation indicates that aeronautical products have evolved to a ‘dominant design’ and entered the ‘specific phase’ of their product life cycle. Innovation in this phase centers on: incremental product improvement, especially for productivity and quality; process technology; technological innovations that offer superior substitutes. The first two of these are aligned with BFC objectives.

The concepts of ‘value’ and ‘best lifecycle value’ are introduced as conceptual frameworks. Value is offered as a metric for BFC. Risk management is intimately tied to achieving value and needs to be integrated into aeronautical processes. The process technology area is addressed by considering ‘lean’ practices for design, engineering and manufacturing. Illustrative results of process improvements from the seven-year Lean Aerospace Initiative research programme at MIT indicate opportunities to achieve BFC. Concluding remarks offer some challenges to industry, government and academics in aeronautical design, engineering and manufacturing.

This paper has arisen from a more general investigation aimed at identifying appropriate manned aircraft that would make suitable stand-off cruise missile platforms. One primary measure of an aircraft's ability to fulfil such a role is found in its payload-range envelope for flight profiles involving one or more payload drop. To this end, a generalisation of the Breguet Range equation is developed initially for a radius-of-action scenario in which the entire payload is ejected at altitude at the designated drop point and the aircraft returns home. The more general problem is then addressed, namely that of establishing the payload-range envelope for an aircraft flying a predefined mission profile comprising a maximum of four consecutive legs each separated by a single payload release point. The ability to include a partial payload release at any drop point is built into the model to permit planners to investigate typical ‘what if?’ effects on the overall mission.

Six numerical examples are included, all based upon the Boeing C-17 as the parent aircraft. These examples describe in detail missions of increasing complexity in order to demonstrate the versatility and efficacy of the current method.

Attention is confined to turbojet-powered aircraft cruising in the stratosphere, although the analogous case of a turboprop-powered aircraft can be simulated simply by amending a single group of constants that appears in all the formulae. Neither air-to-air refuelling policies, nor complicating atmospheric effects such as headwinds, are considered in this paper.