In January 2017, a business jet flew in Norway on a short repositioning flight with two pilots onboard, no passengers or cargo. Initially, the take-off proceeded as normal but as the landing gear was retracted both pilots observed that the airspeed was rapidly approaching the flap limiting speed of 200kts. When the flaps were fully retracted at a height of approximately 2,100ft above ground level, the crew experienced a violent nose-down pitch motion. Control was regained at a height of approximately 170ft above ground level and, following the accident, data from the flight data recorder showed that the aircraft experienced –2.62G during the pitch upset. A tailplane stall due to icing was suspected; however, the flight data recorder, being limited to 36 parameters, was not able to confirm this. For expediency during the accident investigation process, a simplified, linear flight dynamics model was developed using Matlab/Simulink to assess static and dynamic stability for a range of tailplane efficiency factors to simulate the effects of tailplane icing.

The accuracy of several numerical schemes for solving the lifting-line equation is investigated. Circulation is represented on discrete elements using polynomials of varying degree, and a novel scheme is introduced based on a discontinuous representation that permits arbitrary polynomial degrees to be used. Satisfying the Helmholtz theorems at inter-element boundaries penalises the discontinuities in the circulation distribution, which helps ensure the solution converges towards the correct, continuous behaviour as the number of elements increases. It is found that the singular vorticity at the wing tips drives the leading-order error of the solution. With constant panel widths, numerical schemes exhibit suboptimal accuracy irrespective of the basis degree; however, driving the width of the tip panel to zero at a rate faster than the domain average enables improved accuracy to be recovered for the quadratic-strength elements. In all cases considered, higher-order circulation elements exhibit higher accuracy than their lower-order counterparts for the same total degrees of freedom in the solution. It is also found that the discontinuous quadratic elements are more accurate than their continuous counterparts while also being more flexible for geometric representation.

Whirl flutter is an aeroelastic instability that affects aircraft with propellers/rotors. With their long and flexible rotor blades, tiltrotor aircraft are particularly susceptible. Whirl flutter is known to have destroyed aircraft and in the best case it constitutes a fatigue hazard. The complexity of whirl flutter analysis increases significantly with the addition of nonlinearities, due to the more complex dynamical behaviours that emerge as a result. Most whirl flutter stability analyses in current literature are grounded in linear theory, preventing the full discovery of the nonlinearities’ effects. Continuation and bifurcation methods (CBM) may instead be used to fully appreciate and analyse the effects of the presence of nonlinearities. Previous CBM-based work on nonlinear gimballed hub rotor-nacelle models, representing those found on tiltrotor aircraft, are capable of whirl flutter in parametric regions declared safe by linear analysis. Furthermore, it was found that they are capable of complex behaviours including limit cycle oscillations, quasi-periodic behaviour and even chaos, though the whirl flutter implications of such behaviours has not been explored. This paper investigates the impact of a smooth structural nonlinearity on the whirl flutter stability of a basic gimballed rotor-nacelle model, compared to its baseline linear stiffness version. A 9-DoF model with quasi-steady aerodynamics, a flexible wing and blades that can move both cyclically and collectively in both flapping and lead-lag motions, producing gimbal flap-like behaviour, was adopted from existing literature. A smooth stiffness nonlinearity was introduced in the blade flapping stiffness and CBM was used to find the new whirl flutter behaviours created by the presence of the nonlinearity. Time simulations, Poincaré sections and spectral analysis were then used to investigate the various behaviours found. This in turn allowed recommendations to be made concerning preferable and/or hazardous parameter combinations of use to the tiltrotor designer.

This paper presents the results of the experimental study carried out to address the issues of base heating and smooth separation of the stage of launch vehicles. The pressure at the base of a convergent-divergent circular nozzle, from which Mach 1.8 jet emanates, attached to an annular shroud of larger area is controlled by providing air vents on the shroud. On the shroud, vent holes were made at different azimuthal locations, to entrain the surrounding air mass at a higher pressure, pa, to increase the low-pressure, pb, at the base region, caused by the suction creating large-scale vortices formed owing to the sudden expansion of the jet emerging from the nozzle into the shroud. For different number and size of the vents on the shroud, the base pressure was measured. This measurement was done at five levels of overexpansion of the nozzle in the range from –64% to –58%. It is found that increase in vent area results in increase of base pressure, up to some limiting level of the area. Also, the increase of base pressure for the case of vents closer to the nozzle exit is found to be marginally more than the increase caused by vents at distances away from the nozzle exit. Increase of base pressure can be regarded as an advantage not only from base heating point of view but also from the point of view of deflection of the plume to the shroud wall for uniform melting of the pyro layer bonding the stages of the launch vehicle, leading to a smooth separation of the launch vehicle stages.

This paper presents progress towards a transition modelling capability for use in the numerical solution of the Reynolds-averaged Navier-Stokes equations that provides accurate predictions for transonic flows and is thus suitable for use in the design of wings for aircraft flying at transonic speeds. To this end, compressibility corrections are developed and investigated to extend commonly used empirical correlations to transonic flight conditions while retaining their accuracy at low speeds. A compressibility correction for Tollmien-Schlichting instabilities is developed and applied to a smooth local correlation-based transition model and a stationary crossflow instability compressibility correction is included by adding a new crossflow source term function. Two- and three-dimensional transonic transition test cases demonstrate that the Tollmien-Schlichting compressibility correction produces substantially improved agreement with the experimental transition locations, particularly for higher Reynolds number applications where the effects of flow compressibility are expected to be more significant, such as the NASA CRM-NLF wing-body configuration, while the crossflow compressibility correction prevents an inaccurate, upstream transition front. The compressibility corrections and modifications do not significantly affect the numerical behaviour of the model, which provides an efficient alternative to non-local and higher-fidelity approaches, and can be applied to other transport-equation-based transition models with low-speed empirical correlations without affecting their predictive capability in the incompressible regime.

Aerodynamic characterisation from flight testing is an integral subroutine for evaluating a new flight vehicle’s aerodynamic performance, stability and controllability. The estimation of aerodynamic parameters from flight test data has extensively been explored, in the past, using estimation methods such as the equation error method, output error method and filter error method. However, in the current era, non-gradient-based estimation techniques are gaining attention from researchers due to their inherent data-driven optimisation capability to find the global best solution. In this paper, a novel non-gradient-based estimation method is proposed for the aerodynamic characterisation of unmanned aerial vehicles from flight data, which relies on the maximum likelihood method augmented with particle swarm optimisation. Flight data sets of a wing-alone unmanned aerial vehicle are used to demonstrate the capabilities of the proposed method in estimating aerodynamic derivatives. Estimates from the proposed method are corroborated with the wind tunnel test and output error method results. It has been observed that simulated flight vehicle responses using estimated parameters are in good agreement with measured data in most of the manoeuvers considered. Confidence in the estimates of linear and nonlinear aerodynamic parameters is well established with the lower limit of Cramer-Rao bounds, which are minimal. The proposed method also demonstrates good predictability of the quasi-steady stall aerodynamic model by estimating stall characteristic parameters such as aerofoil static stall characteristics parameter, hysteresis time constant and breakpoint. The overall performance of the proposed estimation method is on par with the output error method and is validated with the proof-of-match exercise.

With the aviation industry facing increasing environmental and energy challenges, there has been a growing demand for sustainable aviation fuel (SAF). Previous studies have shown the role of ice accretion, release and blockage in aviation-related incidents and accidents with conventional jet fuel. However, there is no available data that establishes the magnitude of influence new fuel compositions will pose on ice formation and accretion in aircraft fuel systems. A recirculating fuel test rig capable of cooling fuel from ambient to −30°C within 4h was built by Airbus to simulate conditions in an aircraft wing tank and allow characterisation of ice accretion. The key characteristic was the pressure drop across an inline fuel strainer for the different SAF explored but visual analysis of ice accretion on the strainer mesh (filters used in protecting fuel feed pumps) was also performed for individual experimental runs for comparison. Measurements revealed that 100% conventional fuel exhibited a higher propensity to strainer blockage compared to the SAF tested. However, all SAF blends behaved differently as the blending ratio with Jet A-1 fuel had an impact on the pressure differential at different temperatures. Data from this work are essential to establish confidence in the safe operation of future aircraft fuel systems that will potentially be compatible with 100 % SAF.

Flight crews’ capacity to conduct take-off and landing in near zero visibility conditions has been partially addressed by advanced surveillance and cockpit display technology. This capability is yet to be realised within the context of manoeuvring aircraft within airport terminal areas. In this paper the performance and workload benefits of user-centre designed visual and haptic taxi navigational cues, presented via a head-up display (HUD) and active sidestick, respectively, were evaluated in simulated taxiing trials by 12 professional pilots. In addition, the trials sought to examine pilot acceptance of side stick nose wheel steering. The HUD navigational cues demonstrated a significant task-specific benefit by reducing centreline deviation during turns and the frequency of major taxiway deviations. In parallel, the visual cues reduced self-report workload. Pilot’s appraisal of nose wheel steering by sidestick was positive, and active sidestick cues increased confidence in the multimodal guidance construct. The study presents the first examination of how a multimodal display, combining visual and haptic cues, could support the safety and efficiency in which pilots are able to conduct a taxi navigation task in low-visibility conditions.

The ability to perform air-to-air refuelling (AAR) can dramatically extend the utility of helicopters, through effectively providing unlimited range. For helicopters, AAR is typically performed utilising the probe-and-drogue aerial refuelling method. This is a complex manoeuver, where normally both the helicopter and tanker aircraft are operating at the limits of their flight envelopes. In addition, the wake flow from the tanker aircraft can cause a significant disturbance on the refuelling helicopter. This paper presents the initial evaluation of an AAR scenario constructed within DLR’s flight simulator, the Air Vehicle Simulator (AVES), based on current procedures and pilot interviews. A mission task was defined to assess the scenario in AVES and results are subsequently discussed. For pilots unfamiliar to formation flight or HAAR, the results show the difficulty of the flying task itself at the given cueing. Measures for improvement in future investigations are suggested.

For financial and operational reasons many aircraft manufacturers are working on the development of single-pilot commercial aircraft. It is suggested that cargo operations may commence in the early 2030s followed by passenger flights later that decade. Two technological approaches for the development of single-pilot airliners are being developed either based upon extant technology and operating concepts derived from uninhabited aviation systems and military aircraft, or alternatively based upon high levels of onboard autonomy/automation. This review considers the economic, technological, regulatory (safety) and societal acceptance of the single-pilot airliner, and examines some of the operational challenges that airlines may face. It is suggested that while the technological and safety challenges may be resolved, it is the operational challenges that may determine if the concept is ultimately viable.

From arsenal delivery to rescue missions, unmanned aerial vehicles (UAVs) are playing a crucial role in various fields, which brings the need for continuous evolution of system identification techniques to develop sophisticated mathematical models for effective flight control. In this paper, a novel parameter estimation technique based on filter error method (FEM) augmented with particle swarm optimisation (PSO) is developed and implemented to estimate the longitudinal and lateral-directional aerodynamic, stability and control derivatives of fixed-wing UAVs. The FEM used in the estimation technique is based on the steady-state extended Kalman filter, where the maximum likelihood cost function is minimised separately using a randomised solution search algorithm, PSO and the proposed method is termed FEM-PSO. A sufficient number of compatible flight data sets were generated using two cropped delta wing UAVs, namely CDFP and CDRW, which are used to analyse the applicability of the proposed estimation method. A comparison has been made between the parameter estimates obtained using the proposed method and the computationally intensive conventional FEM. It is observed that most of the FEM-PSO estimates are consistent with wind tunnel and conventional FEM estimates. It is also noticed that estimates of crucial aerodynamic derivatives ${C_{{L_\alpha }}},\;{C_{{m_\alpha }}},\;{C_{{Y_\beta }}},\;{C_{{l_\beta }}}$ and ${C_{{n_\beta }}}$ obtained using FEM-PSO are having relative offsets of 2.5%, 1.5%, 6.5%, 3.4% and 7.6% w.r.t. wind tunnel values for CDFP, and 1.4%, 1.9%, 0.1%, 9.6% and 7.5% w.r.t. wind tunnel values for CDRW. Despite having slightly higher Cramer-Rao Lower Bounds of estimated aerodynamic derivatives using the FEM-PSO method, the simulated responses have a relative error of less than 0.10% w.r.t. measured flight data. A proof-of-match exercise is also conducted to ascertain the efficacy of the estimates obtained using the proposed method. The degree of effectiveness of the FEM-PSO method is comparable with conventional FEM.

The need for innovative solutions to enable aerial platforms to fly faster, higher, and longer continues to remain a primary focus for airframe designers. This paper outlines work undertaken to apply a morphing wing warping technology onto a generic Unmanned Aerial Vehicle to deliver enhanced flight performance, efficiency and control capabilities. The prototype employs wings of novel construction which provide both near resistance-free compliance in twist as well as adequate structural stiffness to resist applied loads; all while preserving an aerodynamically smooth surface. Used in combination with developed and integrated closed-loop feedback control architecture, a real-time, non-linear, span-wise wing twist adjustment capability required for optimised flight under differing operating conditions and flight requirements, is demonstrated. Experimental results obtained from a wind tunnel test program show up to a 72% increase in lift to drag ratio under certain conditions compared to a fixed baseline providing some confidence that the combination could be used to realise a step change in flight performance.

We review the development of the contra-rotating propellers from the origins to the present. Initially, these systems were proposed to increase speed, then to increase propulsive efficiency, and thus reduce fuel burn. Ultimately, they hit another environmental limit: too much noise. Acoustics has been in fact the main focus of the development in the past 30 years. Pioneering work done across countries demonstrated several unique features of this propulsor. Various embodiments are available, namely contra-rotating, counter-rotating, co-axial, tandem, open rotor and prop-fans, collectively named contra-rotating propellers. This review only considers concepts that have been applied to real aircraft, prototypes that are known to have been flight tested (about 70 vehicles), or representative laboratory models. Five classifications are proposed: pioneers (before 1940), golden years (1940–1950), Western airplanes (1950s onwards), Soviet-Russian airplanes (1950s onwards) and modern developments (1980s onwards). Selected experimental aircraft and laboratory concepts are mentioned, where these appear to advance the state-of-the-art. Power plants evolved from internal combustion engines to the modern gas turbine engines requiring new solutions. Engine layouts and propulsion configurations are analysed where appropriate. It is concluded that propulsive efficiency can only be achieved at a cost of multiple engineering problems, some of which remain unsolved.

For a class of uncertain systems with large-error sensing, the low-order stable signal corrector and observer are presented for signal correction and uncertainty estimation according to completely decoupling estimation. The model-free signal corrector can reject the bounded stochastic disturbance/error in global position sensing, and system uncertainty can be estimated by the observer, even the existence of large disturbance in position sensing. Furthermore, a general form of signal corrector is given. The describing function method is used to analyse the robustness of the corrector in frequency domain, and the parameter selection rules are presented. The merits of the signal corrector includes its model free, gain-bounded stable structure, sufficient rejection of bounded stochastic disturbance/error in sensing and ease of parameters’ regulation. The corrector and observer are applied to a UAV navigation and control for large disturbance/error corrections in position/attitude angle and the uncertainties estimation in the UAV flight dynamics. The control laws are designed according to the correction-estimation results. Finally, experiments demonstrate the effectiveness of the proposed method.

Problem: Why did the pilots do that? Human error is a reasonably common retrospective assignment of responsibility tied to undesirable aeronautical safety occurrences. Although retributive justice has long been accepted in aviation, its effectiveness in preventing recurrence is minimal. Airmen tend to decide based on their best knowledge with the available resources in intrinsically fallible systems in the ultra-safe high-risk aviation industry.

Method and Results: This paper sheds light on Safety as Capacity under the vanguardist Safety-II perspective and examines procedures as static tools incapable of sustaining safety. It discusses the prejudice in non-critical adherence to procedural compliance beyond creating bureaucratic work environments permissible to sanction workers against regulations. Disputing the safety gain in a retrospective analysis of mishaps, the paper instils the airmen as solution elements to sustain safety at the management of context, a fundamental aspect of Safety-II.

Impact on Industry: A systemic deficiency in civilian pilot training is exposed, and an independent organisational Safety Capacity assessment tool to air operations is provided. The main debate is the synergetic interaction between aircrew’s aeronautical decision-making skills and organisational Safety-II as safety capacity. The pilot’s preparedness to analyse, create and evaluate outside forecasted protocols in modern aviation environments is discussed. These dynamics are revised in their inter-reliability known as Safety as Capacity.

This paper investigates the hydrodynamic near-field of a NACA 16-616 aerofoil over a range of angles-of-attack, encompassing the pre-stall, stall and post-stall flow regimes. In both the static pressure and the pressure fluctuation results, it is shown that each flow regime is easily distinguished, and it is further shown that each regime has different spectral behaviour and boundary layer characteristics. It is found that the NACA 16-616 aerofoil stalls by an abrupt leading-edge mechanism, characterised by a sudden change in the static pressure and unsteady surface pressure spectra between $16^\circ $ and $17^\circ $ angles-of-attack, but of more interest is that there is a secondary yet significant trailing-edge flow separation mechanism occurring upstream of the trailing-edge and moving further upstream as the angle-of-attack increases in the pre-stall regime. A comparison is made between the spectra and coherence of the unsteady surface pressure of the NACA 16-616 aerofoil and those of the classic NACA 0012 aerofoil and shows that such a secondary mechanism has a significant impact for large pre-stall angles-of-attack on the unsteady surface pressure. This will have a significant impact on the radiated far-field sound, distinguishing the NACA 16-616 aerofoil from aerofoils that do not have this secondary mechanism. The existence and extent of this secondary trailing-edge separation mechanism is further shown by the hot-wire anemometry boundary layer velocity results that indicate separation within the pre-stall regime.

The evaluation of the on-ground casualty risk assessments due to a controlled or uncontrolled re-entry is highly sensitive to the accurate prediction of fragmentation events during an atmospheric re-entry. The main objective of this study is an investigation into the use of peridynamics (PD) to improve the analysis of fragmentation during atmospheric re-entry with respect to currently adopted semi-empirical approaches. The high temperatures characterising such scenarios may substantially impact fragmentation, which requires appropriate modelling of the damage process within the PD method. The damage models in PD require experimentally determined fracture mechanical properties that are unavailable as a function of temperature. This work proposes a numerical methodology to estimate the PD damage parameters changes with temperature to enable the study of fragmentation during atmospheric re-entry. Initially, tensile-testing simulation experiments are performed in peridynamics to calibrate material parameters for steel and aluminium alloys as a function of temperature. Then, a parametric study is carried out to evaluate the temperature-dependent damage model parameters for the same materials. The applicability of the proposed methodology is showcased using a re-entry test case scenario.

Prediction of stall before it occurs, or detection of stall is crucial for smooth and lasting operation of fans and compressors. In order to predict the stall, it is necessary to distinguish the operational and stall regions based on certain parameters. Also, it is important to observe the variation of those parameters as the fan transitions towards stall. Experiments were performed on a contra-rotating fan setup under clean inflow conditions, and unsteady pressure data were recorded using seven high-response sensors circumferentially arranged on the casing, near the first rotor leading edge. Windowed Fourier analysis was performed on the pressure data, to identify different regions, as the fan transits from the operational to stall region. Four statistical parameters were identified to characterise the pressure data and reduce the number of data points. K-means clustering was used on these four parameters to algorithmically mark different regions of operation. Results obtained from both the analyses are in agreement with each other, and three distinct regions have been identified. Between the no-activity and stall regions, there is a transition region that spans for a short duration of time characterised by intermittent variation of abstract parameters and excitations of Fourier frequencies. The results were validated with five datasets obtained from similar experiments at different times. All five experiments showed similar trends. Neural Network models were trained on the clustered data to predict the operating region of the machine. These models can be used to develop control systems that can prevent the stalling of the machine.

Civil aircraft that fly long ranges consume a large fraction of civil aviation fuel, injecting an important amount of aviation carbon into the atmosphere. Decarbonising solutions must consider this sector. A philosophical-analytical feasibility of an airliner family to assist in the elimination of carbon dioxide emissions from civil aviation is proposed. It comprises four models based on the integration of the body of a large two-deck airliner with the engines, wings and flight surfaces of a long-range twin widebody jet. The objective of the investigation presented here is to evaluate the impact of liquid hydrogen tank technology in terms of gravimetric efficiency. A range of hydrogen storage gravimetric efficiencies was evaluated; from a pessimistic value of 0.30 to a futuristic value of 0.85. This parameter has a profound influence on the overall fuel system weight and an impact on the integrated performance. The resulting impact is relatively small for the short-range aircraft; it increases with range and is important for the longer-range aircraft. For shorter-range aircraft variants, the tanks needed to store the hydrogen are relatively small, so the impact of tank weight is not significant. Longer range aircraft are weight constrained and the influence of tank weight is important. In the case of the longest range, the deliverable distance increases from slightly over 4,000 nautical miles, with a gravimetric efficiency of 0.3, to nearly 7,000 with a gravimetric efficiency of 0.85.