With the wide application of quadrotor unmanned aerial vehicles (UAVs), the requirements for their safety and reliability are becoming increasingly stringent. In this paper, based on the feedback of airframe performance health perception information and the predictive function control strategy, the autonomous maintenance of a quadrotor UAV with multi-actuator degradation is realised. Autonomous maintenance architecture is constructed by the predictive maintenance (PdM) idea and the Laguerre function model predictive pontrol (LF-MPC) strategy. Using the two-stage Kalman filter (TSKF) method, based on the established UAV degradation model, the aircraft state and actuator degradation state are predicted simultaneously. For the predictive perception of system health, on the one hand, the system health degree (HD) based on Mahalanobis distance is defined by the degree of airframe state deviation from the expected state, and then the failure threshold of the UAV is obtained. On the other hand, according to the degradation state of each actuator, a comprehensive degradation variable fused with different weight coefficients of multiple actuators degradation is used to obtain the probability density function (PDF) of remaining useful life (RUL) prediction. For the autonomous maintenance of system health, the LF-MPC weight matrixes are adjusted adaptively in real-time based on the HD evaluation, to achieve a compromise balance between UAV performance and control effect, and greatly extend the working time of UAV. Simulation results verified the effectiveness of the proposed method.

To meet the increasing communication demands, the satellites need to be equipped with the high-accuracy and large-aperture antennas. One of the effective methods to construct the modular antennas with ultra-high accuracy and ultra-large aperture is on-orbit assembly technology. During the on-orbit assembly missions, the assembly error is a key factor to affect the surface accuracy of the modular antennas. This paper studies the node design of each module and the assembly error analysis of the modular antennas. A design method of the module nodes is presented with consideration of the assembly gap between two modules. Meanwhile, a soft connection mechanism is designed to ensure the mobility among the assembly modules. To investigate the transmission law of the assembly errors, an analytical model of assembly error is derived based on the exponential product method. In order to establish the deformation surface with rotation and displacement assembly errors, an error ball concept is proposed by the analytical model. To decrease the assembly errors, the actuators are installed among some modules. Moreover, an adjustment method is proposed to obtain the adjustment amounts of actuators. Finally, the correctness of analytical model and the effectiveness of the adjustment method are demonstrated by the numerical simulations.

Design/Build/Fly competitions are attracting increased interest in the training of aerospace engineers at academic level worldwide. These competitions entail fundamental activities in aircraft design, optimization and manufacturing which foster student knowledge not possible in classical academic activities. Over the years, the competitiveness of these contests has increased due to the ever-increasing performance that the aircraft exhibit in the flight event. Mass prediction models, specific for competitions such as Air Cargo Challenge (ACC), are presented in this paper. These models are divided into two development methods: statistical and structure-based equations.

The statistical mass models are developed based on data collected from past ACC editions where model accuracy is mainly dependent on the amount of data available. Three models are derived, one containing all available aircraft and two more obtained by dividing the aircraft into balsa- or composite-dominated structures.

Using the structure-based equations method, where the amount of material required to withstand the stresses that the airplane is subjected to is determined, a model is developed for each one of the three considered wing structural concepts, namely two-cell Carbon-Fibre-Reinforced Plastic (CFRP), CFRP D-box and CFRP tube spar. The tail boom component equation is created independently, while the remaining components masses are determined from coefficients based on geometric characteristics and the computed wing or total masses. The average error associated with these models is inferior to 2% for the total mass.

The results obtained from the application to the considered study cases are also presented, and the validity, accuracy, and application in terms of the design phase for each method are discussed.

The purpose of this paper is to group the flight data phases based on the sensor readings that are most distinctive and to create a representation of the higher-dimensional input space as a two-dimensional cluster map. The research design includes a self-organising map framework that provides spatially organised representations of flight signal features and abstractions. Flight data are mapped on a topology-preserving organisation that describes the similarity of their content. The findings reveal that there is a significant correlation between monitored flight data signals and given flight data phases. In addition, the clusters of flight regimes can be determined and observed on the maps. This suggests that further flight data processing schemes can use the same data marking and mapping themes regarding flight phases when working on a regime basis. The contribution of the research is the grouping of real data flows produced by in-flight sensors for aircraft monitoring purposes, thus visualising the evolution of the signal monitored on a real aircraft.

Values of the pressure coefficient Cp between +1 and -1 can be measured directly in incompressible flow by comparison with the known pressures on a circular cylinder placed in the tunnel working section. The development of the method and the apparatus required are described. The method gives more consistent and accurate readings than can easily be obtained with normal manometer measurements, particularly when the pressure differences are small and the tunnel speed not very steady. Since the coefficients depend on Reynolds number only, less precise measurement and control of tunnel speed are required, unless, the coefficients to be measured are subject to critical scale effects. Possible ways of extending the range of coefficients which can be measured are suggested.

It is shown that a pair of supersonic inviscid conical flow fields can exist on either side of a conical deflagration wave. The configuration is relevant to the base flow question and indicates how base drag may be alleviated by burning. Results of exact computations are presented as well as those derived from a slender-body theory.

The solution has been obtained for the problem of a uniformly compressed and symmetrically loaded circular ring plate of linearly varying thickness. Four particular cases have been discussed and numerical values of the maximum deflection have been obtained for various sizes of the hole of the ring.

Detailed measurements have been made of the boundary-layer development on a small two-dimensional aerofoil with supercritical flow and a weak shock wave, together with similar measurements on the tunnel side wall opposite the aerofoil surface. The Reynolds number of the test is similar to that found in the turbines of jet engines and there is a strong favourable pressure gradient ahead of the interaction of the shock with the boundary layer as often occurs in turbine blade passages. However, whereas the boundary layer on the aerofoil is thin and of the same thickness as that on a turbine blade, the thicker boundary layer on the wall is more typical of that on the hub or casing. The experimental results are compared with results from a wide range of calculation methods. One interesting conclusion from these comparisons is the fact that prediction methods which perform well for the thin boundary layers on the aerofoil do not necessarily perform as well for the thicker boundary layers on the wall.

Expressions are obtained for the radial, tangential, and shear stresses in an infinite plate containing a variably reinforced circular hole. The variation in the cross-sectional area of the reinforcement is chosen so that the maximum tensile resistance of the reinforcement occurs at points around the hole where the stresses in the plate tend to be high. It is shown that when the reinforcement has a given weight, the effect of varying its cross section is to reduce the stress concentration in the plate.

The following particular applied stress systems are considered:—

1. (i) Uniform tension or compression in one direction.

2. (ii) Uniform shear.

3. (iii) Unequal principal stresses.

For the particular cases of uniform end loading and unequal applied stresses in the ratio 2:1, data have been obtained which enable the designer to determine the stress distribution in plates with reinforcements having a reasonably wide range of practical dimensions.

The present position relating to the effect of repeated loads on aircraft structures is reviewed from the point of view of the capacity of aircraft structures to withstand such loads and also of their frequencies of occurrence in the life of the aircraft.

A rough account is given of the statistical methods by which such frequencies can be estimated, and the conclusions to which the use of these methods leads are discussed.

The note ends with an outline of a suggested programme of research and a brief statement of the conclusions reached.

Propeller synchrophasing control is an active method to reduce the noise and vibration of turboprop aircraft without additional weight and power. Phase control accuracy has a great influence on the noise reduction effect of synchrophasing. An integrated power/speed/synchrophasing control strategy is proposed to improve the control precision. Speed/phase control transformation logic based on a multi-blade phase plane is adopted which can take both the rapidity of speed response and phase control precision into account, but there exists switching oscillation during the mode transform process. In order to suppress the phase fluctuation due to exterior disturbance, a slave-slave control scheme is provided to take place of a master-slave scheme. Simulation results based on an integrated turboprop engine/propeller real-time non-linear model show that speed/phase integration logic can improve the response rapidity of both the speed and phase. The precision of the control system is verified to be in acceptable range.

In this paper, structural modelling and dynamic analysis methods reflecting the characteristics of a liquid propellant were developed for a pogo analysis. The pogo phenomenon results from the complex interaction between the vehicle structural vibration in the longitudinal direction and the propulsion system. Thus, for an accurate vibration analysis of a liquid propellant launch vehicle, both the consumption of the liquid propellant and the change in the stiffness reflecting the nonlinear hydroelastic effect were simultaneously considered. A complete vehicle structure, including the liquid propellant tanks, was analytically modelled while focusing on pogo. In addition, a feasible liquid propellant tank modelling method was established to obtain an one-dimensional complete vehicle model. With these methods, comparative studies of the hydroelastic effect were conducted. Evaluations of the dynamic analysis of a reference vehicle were also conducted during the first burning stage. The numerical results obtained with the present orthotropic model and the dynamic analysis method were found to be in good agreement with the natural vibration characteristics according to previous analyses and experiments. Additionally, the reference vehicle showed the estimated occurrence of pogo in the first structural mode when compared with the frequencies of the propellant feeding system. In conclusion, the present structural modelling and modal analysis procedures can be effectively used to analyse dynamic characteristics of liquid propellant launch vehicles. These techniques are also capable of identifying the occurrence of pogo and providing design criteria related to pogo instability.

The strip-theory equations of flow through a ducted fan were the basis of a method of designing a fan for a given performance. The merit of the method was that it amounted to an explicit solution of the design problem, no iteration being necessary.

In the inverse problem, of determining the flow past a given fan in conditions other than those of the design, two relations—the motor or tunnel characteristics, for example—will be given between the four quantities P, Ω, η, u (the motor power and angular velocity, the fan efficiency and the air speed at the fan). The problem is to determine the values of these four quantities together with details of the flow past any blade element.

This paper shows that to a high degree of accuracy the performance of a ducted fan can be summed up by saying that P/Ω3 and η are both functions of Ω/u only, functions which depend only on the fan shape. Solution by graphical means is therefore almost instantaneous.

The fan-tunnel combination is shown to be governed by extremely simple equations which amount to saying that the tunnel power factor is a function of Ω/u only which, again, can be calculated once for all for a given fan and tunnel.

The paper considers the distribution of downwash at the surface of rectangular planforms with prescribed subsonic aerodynamic loading. Three separate aspects of the problem are treated analytically, the results of each being tested against those derived from an accurate numerical procedure (Section 2). Asymptotic expressions for large and small aspect ratio are formulated and shown to apply over a wide range of aspect ratio (Section 3). Downwash routines from certain existing lifting-surface methods are studied, and their patterns of convergence are illustrated and compared (Section 4). A logarithmic singularity near the leading tip corner is identified, but it is observed that this can exist without serious detriment to the lifting-surface methods (Section 5).

A recent calculation method proposed by Curle has been applied to the flow U=U0ξe-ξ for which a computer solution exists. An earlier approximate method due to Head, presented here in a simplified form, has been applied to the same problem. All three solutions are found to be in close agreement. A further problem, already examined by Curle, is treated by Head’s method. Again the results are in good agreement with each other and with an exact solution obtained subsequently.

This paper gives some identities concerning the flexibility of certain elastic structures, including flat plates, at the onset of buckling and subsequently. The identities have a useful application in determining the flexibility and the rate of change of flexibility with load for structures in which the mode in the buckled state changes continuously with load.

In Reference 3, attention is drawn to the difficulties of measuring the streamwise extent of the roughness from insects. The present paper deals with the problem theoretically for an aerofoil in two-dimensional, incompressible flow. A tentative approach to the determination of effective excrescence height downstream of the leading-edge zone is also advanced. The application of these investigations, in conjunction with the analysis from Ref. 3 regarding the critical conditions for premature transition, leads to estimates of the amount of significant roughness which are in good agreement with flight observation.

In order to increase the speed, precision and robustness against the engine failure in solving optimal endo-atmospheric ascent trajectory of a launch vehicle, a rapid multi-layer solving method with improved numerical algorithms was proposed. The proposed method is capable of decomposing a large number of intervals into multiple layers with advantageous convergence property. Firstly, the problem of solving optimal endo-atmospheric ascent trajectory, which was subjected to path constraints and terminal constraints, was transformed into a Hamilton Two Point Boundary Value Problem (TPBVP). Then, through the finite difference method and numerical solving algorithm, the Hamilton TPBVP was iteratively solved with fewer initial discrete intervals. The initial values of higher-layer iterations were obtained by interpolating convergent solutions at sparse nodes into the doubly discrete nodes of high layers. The process was repeatedly performed until the solving precision met the requirements. To decrease the calculation load in solving TPBVPs, two improved solving algorithms without and with fewer Jacobian calculations were studied, respectively the Derivative-free Spectral Algorithm for Nonlinear Equations(DF-SANE) combined with the improved derivative-free nonmonotone line search strategy, and the Modified Newton method with a relaxation factor in combination with the Inverse Broyden Quasi-Newton method, denoted as ‘MN-IBQ’. Simulation verifications showed that the multi-layer method had significantly higher solving speed than the single-layer method. For the improved numerical algorithms, the DF-SANE was trapped in the local convergence problem. While using the proposed MN-IBQ can further increase the solving rate. Typical engine failure simulations showed that the multi-layer method with the MN-IBQ algorithm had not only significantly higher solving speed but also stronger robustness, where the traditional single-layer method could not adapt. In addition, the thrust loss tolerance limits for the multi-layer solving method were given for different engine failure times. The results show promising potential of the proposed approach in trajectory online generation and closed-loop guidance of launch vehicles at the endo-atmospheric ascent stage.

Among the many problems which face airframe manufacturers, cooling is certainly one of the most important, particularly as a special watch must be kept on the type of solution to be adopted, since the lowest fuel penalty must be incurred in flight.

At the Aerodynamische Versuchsanstalt Gottingen (AVA), base pressure measurements were made on five elliptic cones. The ratios of the axes of the ellipses were: 1:12, 1:3, 1:1, 3:1, and 12:1. All the cones had the same volume and the same length. The investigations were carried out for Mach numbers M∞=1·50, 1·73, and 1·98 at angles of incidence between about —2 degrees and about 8 degrees. Reynolds numbers, based on a mean length, lm, varied from 2·5×106 to 3·0×106; the boundary layer approaching the base was turbulent. The base pressure coefficients are given as a function of geometry. Some of the results were compared with the base-flow theory of Korst and reasonable agreement was found.