Spinning/Rolling Airframe Missiles (RAM) mostly use an analog (ON-OFF type) control approach that deflects control surfaces (fins) at minimum and maximum positions continuously, and control is achieved by applying phase shift between the minimum and maximum deflections during a complete roll cycle. Therefore, the control signal shape changes continuously during the roll cycle. In this study, a novel single channel digital controller is designed and tested for a high spinning rate (10–20 Hz) RAM. The digital controller adjusts the amplitude of fin deflections instead of applying a phase shift. In this way, control signal shapes are predetermined in design to completely decouple yaw and pitch dynamics. At the beginning of each roll cycle, the algorithm decides on the control signal shape and amplitude to apply throughout the cycle. Delays on the actuator system and sensor measurements which might lead to instability at high spinning rates are handled effectively thanks to the predetermined control signal shapes that are changing with 90-degree intervals. Detailed geometry of a surface-to-air spinning missile is used to obtain aerodynamic coefficients for the entire flight regime (i.e. from launch to terminal phase) via Missile DATCOM. The 6-DOF flight dynamics model and the controller algorithms are built in MATLAB/Simulink environment. The proposed digital controller is tested systematically for various scenarios and the performance is compared with the conventional analog control approach. The digital controller gives better performance compared to the analog approach under the influence of servo delays and sensor noise.

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.

This paper describes a new efficient method for the construction of an approximately balanced aerodynamic Reduced Order Model (ROM) via the frequency domain using Computational Fluid Dynamics data. Time domain ROM construction requires CFD data, which is obtained from the DLR TAU RANS or Euler Linearised Frequency Domain (LFD) solver. The ROMs produced with this approach, using a small number of frequency simulations, are shown to exhibit a strong ability to reconstruct the system response for inviscid flow about the NLR7301 aerofoil and the FFAST wing; and viscous flow about the NASA Common Research Model. The latter demonstrates that the reduced order model approach can reconstruct the full order frequency response of a viscous aircraft configuration with excellent accuracy using a strip wise approach. The time domain models are built using the frequency domain, but also give promising results when applied to reconstruct non-periodic motions. Results are compared to time domain simulations, showing good agreement even with small ROM sizes, but with a substantial reduction in calculation time. The main advantage of the current model order reduction approach is that the method does not require the formation and storage of large matrices, such as in POD approaches.

For a common micro-satellite, orbiting in a circular sun-synchronous orbit (SSO) at an altitude between 500 and 600km, the satellite attitude during off-nadir imaging and staring-imaging operations can be up to ±45 degree on roll and pitch angles. During these off-nadir pointing for both multi-trip operation and staring imaging operations, the spacecraft body is commonly subject to high-rate motion. This posts challenges for a spacecraft attitude determination subsystem called Gyro Stellar Inertial Attitude Estimate (GS IAE), which employs gyros and star sensors to maintain the required attitude knowledge, since star trackers will severely degrade attitude estimation accuracies when the spacecraft is subject to high-rate motion. This paper analyses the star motion-induced errors for a typical star tracker, models the star motion-induced errors to assess the performance impact on the attitude estimation accuracy, and investigates the adaptive extended Kalman filter design in the GS IAE while evaluating its effectiveness.

The high operational cost of aircraft, limited availability of air space, and strict safety regulations make training of fighter pilots increasingly challenging. By integrating Live, Virtual, and Constructive simulation resources, efficiency and effectiveness can be improved. In particular, if constructive simulations, which provide synthetic agents operating synthetic vehicles, were used to a higher degree, complex training scenarios could be realised at low cost, the need for support personnel could be reduced, and training availability could be improved. In this work, inspired by the recent improvements of techniques for artificial intelligence, we take a user perspective and investigate how intelligent, learning agents could help build future training systems. Through a domain analysis, a user study, and practical experiments, we identify important agent capabilities and characteristics, and then discuss design approaches and solution concepts for training systems to utilise learning agents for improved training value.

The design of a two-dimensional supersonic inlet with large bleed window at low Mach number was developed. Numerical simulation and wind tunnel experiments were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the single-degree-of-freedom variable geometry scheme adopted in this paper guarantees the steady work of the inlet over a wide speed range. The large bleed window caused by rotation of the compression ramp appears near the throat at low Mach number. Low-pressure airflow near the bleed window neutralises the original high-pressure airflow behind the shock train, which decreases the overall pressure of the downstream region of the internal contraction section. To match the lower pressure, the structure of the shock train changes from strong $\lambda$-type to weak $\lambda$-type, and finally to a normal shock wave as backpressure increases at Mach number 2.5. Herein, the total pressure recovery coefficient of the inlet near the critical condition improves by 8.5% as the backpressure ratio (Pe/P0) adds from 13 to 14.6 at Mach number 2.5. It proves that the scheme is effective on terminal shock wave control and inlet performance improvement. In addition, due to the background wave and the bleed window, two kinds of shock wave oscillation occur when the backpressure ratio is 13.1.

This study is focused on the development of longitudinal aerodynamic models for steady flight conditions. While several commercial solvers are available for this type of work, we seek to evaluate the accuracy of an open source software. This study aims to verify and demonstrate the accuracy of the OpenFoam solver when it is used on basic computers (32–64GB of RAM and eight cores). A new methodology was developed to show how an aerodynamic model of an aircraft could be designed using OpenFoam software. The mesh and the simulations were designed only using OpenFoam utilities, such as blockMesh, snappyHexMesh, simpleFoam and rhoSimpleFoam. For the methodology illustration, the process was applied to the Bombardier CRJ700 aircraft and simulations were performed for its flight envelope, up to M0.79. Forces and moments obtained with the OpenFoam model were compared with an accurate flight data source (level D flight simulator). Excellent results in data agreement were obtained with a maximum absolute error of 0.0026 for the drag coefficient, thus validating a high-fidelity aerodynamic model for the Bombardier CRJ-700 aircraft.

This study proposes a new operational concept of the Point Merge System, called Multi-Arrival Route Point Merge System (MAR-PMS), which is an air traffic control method used to sequence aircraft arrivals in a given terminal control area. The proposed concept enables the additional arrival routes that have an angular difference to each sequencing leg. Furthermore, a time-indexed 0-1 linear programming model is formulated. The obtained results are validated in a real time simulation. The comparison results of PMS and MAR-PMS show that the average reduction of 19% of total flight time, 23% of total flight distance, and 19% in total fuel burned and reduction in CO2 emissions in favor of a proposed concept.

This paper addresses the potential impact of adverse aircraft-pilot couplings on tiltrotor safety, when a pilot or autopilot attempts to constrain flight dynamics with strong control. The work builds on previously published research on the theory and application of constrained flight to fixed- and rotary-wing aircraft. Tiltrotor aircraft feature characteristics from both types of aircraft and how these determine behaviour in a unique manner is investigated using a FLIGHTLAB simulation model of the XV-15 aircraft. Two different scenarios are explored in detail, using linearised models that reflect the flight-physics of stability for small deviations from trim. First, the control of vertical flight path with longitudinal cyclic pitch and elevator, and the consequences for the stability of the aircraft surge mode and short-period pitch-heave mode. The classical surge-mode instability for flight at speeds below minimum power is shown to apply to the tiltrotor in helicopter mode but alleviated in conversion and airplane modes. The impact on the short–period mode is seen to be a trade-off between the stabilising pitch attitude and destabilising incidence (angle-of-attack) contributions to the flight-path angle. The second example concerns strong control of roll attitude in the presence of adverse aileron-yaw. Here, the yaw-sway motion can be driven unstable, a problem encountered on fixed-wing aircraft with weak weathercock stability, but rare in the rotorcraft world. For both examples, the loss of stability is expressed as the change in sign of effective damping or stiffness stability derivatives. The explanatory theory for these non-oscillatory or low-frequency aircraft-pilot couplings is presented, along with interpretations in terms of handling qualities criteria. The paper also addresses the question of how to translate the findings into a form of aeronautical knowledge useful for the pilot training community.

For a multi-vectored propeller aerostat with actuator faults, this study presents a fault-tolerant tracking control strategy, which includes fault modeling, observer, force estimation and tracking controller. Fault modeling considers the four types of faults of vectored propellers, namely, thrust offset, thrust efficiency loss, vectored angle offset and vectored angle stuck. Actuator faults can be determined from the fault observer, which identifies the thrust offset from the acceleration difference of the faulty aerostat with the ideal model. For tracking positions, a traditional PID controller is constructed with virtual control, compensated with the estimated fault force. The control allocation scheme is proposed to redistribute the available actuators in case faults occur. Simulation results of position tracking prove the effectiveness of the proposed strategy.

A new type of design domain for a sounding rocket is suggested in this study, which is more intuitive, simplified, and conducive to a single-stage sounding rocket development process. Among the various operation parameters, this study identifies several effective variables, which are also among the most practical in a sounding rocket design process. Of the many design variables considered for peak altitude optimisation, arguably the three most effective in shaping the whole system and conferring the most practicality in the design process of a sounding rocket were determined. A simulation-based study was conducted to establish: the effect of the selected parameters on flight performance, and the optimum design conditions of a single-stage sounding rocket in terms of its peak altitude. The simulation result was compared with randomly chosen experimentally tested flight data and validated. With the clear performance curves varying depending on the variables, the combination of the considered design inputs was effective. The new type of design domain and design procedure proposed is expected to constitute a useful reference and deliver practical benefits to the development process for a target-altitude-optimised single-stage sounding rocket.

Arc-jet facilities have been the norm for ablation experiments used to calibrate computational models to date. However, the arc jet has a few major limitations and challenges, including non-uniform enthalpy distribution, non-equilibrium state, change of surface quality during testing and the extent of oxidation, to name but a few. A novel plasma technique for preheating axisymmetric heatshield samples in hypersonic impulse facilities is presented herein. The major aim of this innovative work is to help reduce the large variations of ablation rate predictions, space vehicle materials and missile design/testing, obtain strongly coupled hypersonic boundary layers and achieve lower cost of aerothermodynamics experiments. This present work remains one of the most highly anticipated solutions to maximise payload success and replicate high surface temperatures identical to those experienced by real flight vehicles. This work makes a useful contribution to re-entry studies under conditions that replicate the characteristics of re-entry flights. Future applications for the technique are expected to be found in hypersonic impulse facilities that can simulate the true flow energy under re-entry conditions.

The objective of the present work is to estimate the performance of a turbojet engine during Fluidic Thrust Vectoring (FTV) employed by injecting the secondary-jet at the throat of a convergent nozzle. The nozzle performance maps and effective nozzle throat area obtained from experiments are coupled with the performance of a conventional engine (without FTV) using an iterative algorithm developed as a part of this work. The performance is estimated for different flow rates of secondary-jet sourced either from a separate compressor or the engine’s compressor. During FTV, the operating point shifted towards the surge line with increased turbine entry temperature. The desired and obtained vector angles and thrust magnitudes are different. At high secondary-jet flow rates, the turbine operation moved out of its performance map. These aspects should be incorporated while integrating the FTV at the system level, thus, asserting the importance of FTV studies coupled with engine performance.

In this paper, we expolore Multi-Agent Reinforcement Learning (MARL) methods for unmanned aerial vehicle (UAV) cluster. Considering that the current UAV cluster is still in the program control stage, the fully autonomous and intelligent cooperative combat has not been realised. In order to realise the autonomous planning of the UAV cluster according to the changing environment and cooperate with each other to complete the combat goal, we propose a new MARL framework. It adopts the policy of centralised training with decentralised execution, and uses Actor-Critic network to select the execution action and then to make the corresponding evaluation. The new algorithm makes three key improvements on the basis of Multi-Agent Deep Deterministic Policy Gradient (MADDPG) algorithm. The first is to improve learning framework; it makes the calculated Q value more accurate. The second is to add collision avoidance setting, which can increase the operational safety factor. And the third is to adjust reward mechanism; it can effectively improve the cluster’s cooperative ability. Then the improved MADDPG algorithm is tested by performing two conventional combat missions. The simulation results show that the learning efficiency is obviously improved, and the operational safety factor is further increased compared with the previous algorithm.

The stabilisation and control mechanisms of an Unmanned Aerial System (UAS) must be properly designed to ensure acceptable flight performance. During their operation, these mechanisms are subjected to unknown and random environmental effects, making it imperative that all available information should be taken into consideration during the mechanisms’ design process (e.g. system dynamics, actuators, flight conditions and certain criteria requirements such as phugoid and short modes for longitudinal dynamics, and roll subsidence, spiral and Dutch-roll modes for lateral dynamics) in order to guarantee flight stability. Therefore, this paper introduces a novel methodology for the stabilisation and control of the UAS-S45 Bálaam, designed and manufactured by Hydra Technologies. This methodology uses composite controllers that combine feedback Linear Quadratic Regulators (LQR) and Proportional Integral Feed-Forward (PI-FF) compensation controller for stabilisation and tracking tasks, respectively. Furthermore, a Generalised Extended State Observer was implemented to provide robustness to the closed loop dynamics by introducing disturbance compensation. Furthermore, an Adaptive Neuro-Fuzzy Inference System (ANFIS) was adopted to perform a gain scheduling by computing the gains of each composite controller for certain unknown trim conditions within a given flight domain. Finally, several numerical assessments were performed to highlight the efficiency of the proposed methodology.

Reinforced cables are usually installed on flexible airship structures to enhance their load-bearing capability. However, reinforced cables also increase the total weight of the airship. In order to find a balance between large loading-bear capability and light weight, a multi-objective optimisation scheme based on the genetic algorithm NSGA-II is put forward for the reinforced cable distribution on the airship. Firstly, different cable distribution schemes are presented according to engineering experience and the optimal one is determined by load analysis. Then, the CAE method and optimisation analysis are combined to achieve structure design optimisation. The parametric model of the airship structure with reinforced cables is established by ABAQUS secondary development and the load analysis is carried out. Parameter passing and optimisation algorithm are operated by Isight software and the optimisation analysis is conducted based on the NSGA-II algorithm. Finally, we draw some conclusions of the rules of optimised reinforcing cable distribution. The work of this paper has crucial engineering significance for improving performance of the airship structure design.

In this paper, the prediction of the unsteady flow field over typical high aspect ratio (AR) wings in the transonic flow regime but below the sonic Mach number is of interest. The methodology adopted is a computational approach based on the transonic small disturbance unsteady potential equation. It is shown that the higher AR wings generally have a higher lift coefficient as well as a higher lift-to-drag ratio. With NASA’s common research model (CRM) wing, there is an increase in maximum lift with increasing AR while the induced drag is almost the same. There is also an optimum sweep angle, which is different for each angle-of-attack so that variable sweep lifting surfaces may be designed to provide optimum solutions. The computed flutter speeds indicate an expected reduction with increasing AR.