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In the case of the given design variables and constraint functions, this paper is concerned with the rapid overall parameters design of trajectory, propulsion and aerodynamics for long-range ballistic missiles based on the index of the minimum take-off mass. In contrast to the traditional subsystem independent design, this paper adopts the research idea of the combination of the subsystem independent design and the multisystem integration design. Firstly, the trajectory, propulsion and aerodynamics of the subsystem are separately designed by the engineering design, including the design of the minimum energy trajectory, the computation of propulsion system parameters, and the calculation of aerodynamic coefficient and dynamic derivative of the missile by employing the software of missile DATCOM. Then, the uniform design method is used to simplify the constraint conditions and the design variables through the integration design, and the accurate design of the optimized variables would be accomplished by adopting the uniform particle swarm optimization (PSO) algorithm. Finally, the automation design software is written for the three-stage solid ballistic missile. The take-off mass of 29 850 kg is derived by the subsystem independent design, and 20 constraints are reduced by employing the uniform design on the basis of 29 design variables and 32 constraints, and the take-off mass is dropped by 1 850 kg by applying the combination of the uniform design and PSO. The simulation results demonstrate the effectiveness and feasibility of the proposed hybrid optimization technique.

This paper presents a methodology for aerodynamic optimization of UAV with VTOL capabilities. Aircrafts such as these usually fly at low speeds and due to that low Reynolds numbers are to be expected. The friction drag is highly dependent on the quality of the production process so unless special measures are undertaken, high friction drag coefficients could drastically influence overall performance of the aircraft. Changes of the geometrical parameters influence not only the induced drag of the wing, but also the distribution of the base drag due to sensitivity to changes of the Reynolds numbers. In order to determine the optimal geometrical parameters of the wing, a code for wing performance analysis was written. All necessary factors were calculated by utilizing the Glauert's solution of the Prandtl's equation for multi-segmented wings. By including experimental data of numerous airfoils optimized for low Reynolds numbers, the base drag distribution, along with the induced drag of the wings were calculated for a wide range of angles-of-attack. The obtained results are presented through diagrams and the methodology for the selection of the highest efficiency wing is described. The design of the T - shaped stabilizer was achieved by utilizing analytical methods while the Vortex Lattice Method, DATCOM and CFD were used for verification purposes.

This paper proposes an attitude control method for variable-wing missiles. First, the aerodynamic parameters of the missiles are generated using the Missile DATCOM, and then they are fitted into the polynomial functions for controller design. Due to the morphing structure and strong nonlinearity of variable-wing missiles, working-point-based model linearization will lose model accuracy and consequently degrades the control performance. In this paper, an attitude controller for morphing missiles is proposed by combing the disturbance observer and the feedback linearization (FL) method. Theoretical research and numerical simulations show that the designed control system is suitable for morphing missiles, and has strong robustness to system parameter perturbations and external disturbances.

Unmanned Aerial Gliders (UAGs), are a class fixed-wing Unmanned Aerial Vehicles (UAVs) that are particularly well-suited to applications involving long endurance and range. By relying on atmospheric energy, most commonly in the form of thermal updraft, UAGs are able to gain altitude and extend flight time while preserving on-board battery energy to further prolong the mission. While previous research has investigated autonomous soaring for UAGs, the focus has been on detecting and exploiting thermals while in gliding flight. In this work, soaring algorithms are adapted to detect the presence of thermal updraft while in either powered or gliding flight. This allows the aircraft to latch onto thermals at any point during flight and capitalize on atmospheric energy to preserve on-board battery energy. Furthermore, to generate a sink polar for use in the soaring algorithms, both analytically and experimentally, digital DATCOM is used to analyze the aerodynamics of the aircraft in the former case, and a Rauch-Tung-Striebel smoother is proposed to filter experimental flight data for postprocessing. The algorithm is subsequently integrated into the PX4 flight stack and testing is performed in a Software-in-the-Loop environment. Simulation results show the improvement in efficiency gained by using atmospheric energy through all phases of flight.

The prime focus of this work is to estimate stability and control derivatives of an airship in a completely nonlinear environment. A complete six degrees of freedom airship model has its aerodynamic model as nonlinear functions of angle of attack. Estimating the parameters of aerodynamic model in a nonlinear environment is challenging as it demands an exhaustive dataset that could cover the entire regime of operation of airship. In this work, data generation is achieved by simulating the mathematical model of airship for different trim conditions obtained from continuation analysis. The mathematical model is simulated using predicted parameter values obtained using DATCOM methodology. A modular neural network is then trained using back-propagation and Adam optimisation algorithm for each of the aerodynamic coefficients separately. The estimated nonlinear airship parameters are found to be consistent with the DATCOM parameter values which were used for open-loop simulation. This validates the proposed methodology and could be extended to estimate airship parameters from real flight data.

The morphing aircraft can change different wing shapes or geometries to achieve the optimal flight performance according to various mission scenarios. In this paper, DATCOM is used to calculate aerodynamic parameters based on Firebee UAV morphing aircraft with different wing configurations and analyze aerodynamic characteristics. A novel adaptive wing morphing strategy for morphing aircraft based on reinforcement learning method is proposed. This method can highly meet the demand of keeping optimal performance in multiple flight conditions, and the adaptive wing morphing strategy, three-loop normal load altitude controller and sliding mode velocity controller can together make sure stability of morphing aircraft during morphing process with good tracking performance.

Aerodynamic shape optimization is usually a loop of an optimization model, an optimizer and an evaluation workflow. A new optimizer is proposed and tested for a typical aerodynamic shape optimization of missile control surfaces with computational fluid dynamics (CFD). The new optimizer emphasizes the use of machine learning techniques, reinforcement learning and transfer learning, to improve performance and efficiency. Reinforcement learning is applied to extract the optimization experience from the semi-empirical method DATCOM using deep neural networks. Transfer learning is implemented to reuse the experience as priori knowledge in the CFD-based optimization by sharing neural network parameters. For the considered aerodynamic shape optimization problem of missile control surfaces, a remarkable reduction in the computational time has been accomplished. The new approach significantly decreases the required CFD calls by over 62.5%. Meanwhile, the time spent in the experience extraction and parameter transfer process is negligible.

The rising demand for Unmanned Aerial Systems (UASs) to perform tasks in hostile environments has emphasized the need for their simulation models for the preliminary evaluations of their missions. The efficiency of the UAS model is directly related to its capacity to estimate its flight dynamics with minimum computational resources. The literature describes several techniques to estimate accurate aircraft flight dynamics. Most of them are based on system identification. This paper presents an alternative methodology to obtain complete model of the S4 and S45 unmanned aerial systems. The UAS-S4 and the UAS-S45 models were divided into four sub-models, each corresponding to a specific discipline: aerodynamics, propulsion, mass and inertia, and actuator. The “aerodynamic” sub-model was built using the Fderivatives in-house code, which is an improvement of the classical DATCOM procedure. The “propulsion” sub-model was obtained by coupling a two-stroke engine model based on the ideal Otto cycle and a Blade Element Theory (BET) analysis of the propeller. The “mass and the inertia” sub-model was designed utilizing the Raymer and DATCOM methodologies. A sub-model of an actuator using servomotor characteristics was employed to complete the model. The total model was then checked by validation of each sub-model with numerical and experimental data. The results indicate that the obtained model was accurate and could be used to design a flight simulator. Keywords: Aerodynamics, Aerospace propulsion, Flight dynamics, Unmanned aerial vehicles modeling

This paper focuses on developing an incremental-based model reference adaptive control (MRAC) system for the RKX-200TJ LAPAN Missile. The mathematical model of the missile is represented by six-degree of freedom (6-DOF) derived from the rigid-body dynamic principle. The model was linearized and decoupled into longitudinal plane and lateral- directional plane. This paper reviews flight stability in lateral plane using aerodynamic characteristics and stability derivatives calculated from Digital DATCOM software. The incremental-based MRAC utilizes the Lyapunov function approach to generate adaptive law parameters. The simulation results show that the proposed method can effectively improve the performance of the missile control system when it experiences an external disturbance during flight.

The development of Unmanned Aerial Vehicle (UAV) technology in both developing and developing countries is ac- celerating. In particular, UAV technology innovation for maritime affairs. The Aviation Technology Center has developed several UAVs used for the Maritime Surveillance System (MSS), which are marine surveillance and operations systems based on the use of UAV technology. Until now, several UAV variants have been developed, one of which is being developed is the LSU 05-NG. The main target of the design of the LSU 05-NG is to get a UAV as a lightweight utility platform and can operate in the marine environment and near the sea. Therefore it is necessary to do a static and dynamic analysis to determine the performance of the LSU 05-NG flight. In this paper using the numerical method VLM, CFD and DATCOM to analyze the aerodynamic characteristics of LSU 05-NG. Based on the results of the analysis of the values of CL, CD, CM, Cn, and Cl, it can be concluded that LSU 05-NG has static stability. Because the values CMα 0.

Technological growth had changed the conduction way of the warfare. This work is paying attention with developing the performance of a SSM (surface-to-surface missile) aerodynamically controlled system via both predictive and Neuro-Fuzzy controllers. The analysis and design demand somehow precise system model with different uncertainties via 6-DOF simulation. The executive differential equations of the missile motion are obtained from the model of missile's aerodynamic that built by means of the Missile Datcom software. Hence getting the required aerodynamics stability derivatives by the resulted aerodynamics data, the required transfer functions are calculated based on the equations of motion for the missile. Then the yaw autopilot is designed using the calculated transfer functions and to recognize the command signal produced by the guidance laws. The form of guidance commands are in the lateral acceleration components. By using the whole system model, the computer simulators are made by using the Matlab-Simulink software, where predictive and hybrid AI (Artificial Intelligence) Neuro-Fuzzy Yaw autopilots are compared proving that the more stability and less processing time.

This paper focuses on stability augmentation control system development for fixed-wing Turbojet Unmanned Aerial Vehicle (UAV). UAV model is generated by using digital DATCOM software to calculate aerodynamic derivatives. Longitudinal flight dynamics will be used to synthesize control responses in xb and zb plane. This research of stability augmentation uses Linear Quadratic Regulator (LQR) approach. Flight stability of this UAV comes from a solution to Differential Riccati Equation (DRE) which provides optimal controller gain to adjust the states. The simulation results depict that the proposed method can be considered to improve flight performance in the presence of disturbances.

Autonomous Underwater Vehicle (AUV), which are becoming more and more important in ocean exploitation tasks, needs energy conservation urgently when sailing the complex mission path in long time cruise. As hull lines optimization design becomes the key factor, which closely related with resistance, in AUV preliminary design stage, uncertainty parameters need to be considered seriously. In this research, Myring axial symmetry revolution body with parameterized expression is assumed as AUV hull lines, and its travelling resistance is obtained via modified DATCOM formula. The problems of AUV hull lines design for the minimum travelling resistance with uncertain parameters are studied. Based on reliability-based optimization design technology, Design For Six Sigma (DFSS) for high quality level is conducted, and is proved more reliability for the actual environment disturbance. Keywords: AUV, Uncertainty parameters, RBOD, DFSS

Purpose The purpose of this paper is to present the status of the on-going development of the new computerized environment for aircraft synthesis and integrated optimization methods (CEASIOM) and to compare results of different aerodynamic tools. The concurrent design of aircraft is an extremely interdisciplinary activity incorporating simultaneous consideration of complex, tightly coupled systems, functions and requirements. The design task is to achieve an optimal integration of all components into an efficient, robust and reliable aircraft with high performance that can be manufactured with low technical and financial risks, and has an affordable life-cycle cost. Design/methodology/approach CEASIOM (www.ceasiom.com) is a framework that integrates discipline-specific tools like computer-aided design, mesh generation, computational fluid dynamics (CFD), stability and control analysis and structural analysis, all for the purpose of aircraft conceptual design. Findings A new CEASIOM version is under development within EU Project AGILE (www.agile-project.eu), by adopting the CPACS XML data-format for representation of all design data pertaining to the aircraft under development. Research limitations/implications Results obtained from different methods have been compared and analyzed. Some differences have been observed; however, they are mainly due to the different physical modelizations that are used by each of these methods. Originality/value This paper summarizes the current status of the development of the new CEASIOM software, in particular for the following modules: CPACS file visualizer and editor CPACSupdater (Matlab) Automatic unstructured (Euler) & hybrid (RANS) mesh generation by sumo Multi-fidelity CFD solvers: Digital Datcom (Empirical), Tornado (VLM), Edge-Euler & SU2-Euler, Edge-RANS & SU2-RANS Data fusion tool: aerodynamic coefficients fusion from variable fidelity CFD tools above to compile complete aero-table for flight analysis and simulation.

Purpose This paper aims to present an aerothermodynamic analysis of a new concept of a small hypersonic airplane. Aerodynamics characteristics for different flow conditions encountered during the missions are analyzed. The effects of elevons deflection for pitch control and of the presence of engines on aerodynamic performances are also investigated for different flight conditions. The effects of boundary layer laminar–turbulent transition on aerodynamic heating are studied to preliminarily identify proper materials that can sustain the hypersonic phase. Design/methodology/approach Aerodynamic characteristics are predicted by means of the semi-empirical aerodynamic prediction code Missile DATCOM and computational fluid dynamics simulations. Computational fluid dynamics analysis is also performed to investigate aerodynamic heating phenomenon. Findings Major discrepancies between the results offered by the two methods have been registered in transonic regime, whereas in subsonic and super-hypersonic conditions, Missile DATCOM confirms to be a suitable tool for preliminary design steps. The results of the analysis show that for the identification of the materials that can sustain the hypersonic phase, the turbulent solution must be taken into account. Carbon fiber reinforced ceramics composite materials seem particularly well suited for the nose, wing and vertical tail leasing edges and control surfaces, while titanium alloys could be used for the rest of the vehicle surface. Originality/value This new concept of vehicle is designed both for point-to-point medium range hypersonic transportation and long duration suborbital space tourism missions, by integrating available technologies developed for aeronautical and space systems.

This article presents a structural analysis of the Unmanned Aerial System UAS-S4 ETHECATL. Mass, centre of gravity position and principal mass moment of inertia are numerically determined and further experimentally verified using the ‘pendulum method’. The numerical estimations are computed through Raymer and DATCOM statistical-empirical methods coupled with mechanical calculations. The mass of the UAS-S4 parts are estimated according to their sizes and the UAS-S4 class, by the means of Raymer statistical equations. The UAS-S4 is also decomposed in several simple geometrical figures which centres of gravity are individually computed, weighted and then arithmetically averaged to find the whole UAS-S4 centre of gravity. In the same way, DATCOM equations allows us to estimate the mass moments of inertia of each UAS-S4 parts that are finally sum up according to the Huygens-Steiner theorem for computing the principal moment of inertia of the whole UAS-S4. The mass of de UAS-S4 is experimentally determined with two scales. Its centre of gravity coordinates and its mass moment of inertia are found using the pendulum method. A bifilar torsion-type pendulum methodology is used for the vertical axis(14)and a simple pendulum methodology is used for the longitudinal and transversal axes(12). The test object is installed on a pendulum (simple or bifilar torsion pendulum) which is led to oscillate freely while recording the oscillation's angles and speed, by the means of three sensors (an accelerometer, a gyroscope and a magnetometer) that the calibration is also discussed. Simultaneously, nonlinear dynamic models are developed for the rotational motion of pendulums, including the effects of large-angle oscillations, aerodynamic drag, viscous damping and additional momentum of air. ‘Algorithms of minimization’ are then used to simulate and actualise the dynamic models and finally chose the model that simulated data best fit the experimentally recorded one. Pendulum parameters, such as mass moment of inertia, are lastly extracted from the chosen model. To determine the accuracy of the nonlinear dynamics approach of the pendulum method, the experimental results for an object of uniform density for which the mass moments of inertia are computed numerically from geometrical data are presented along with the experimental results obtained for the UAS-S4 ETHECATL. For the uniform density object, the experimental method gives, with respect to the numerical results, an error of 4.4% for the mass moment of inertia around theZaxis and 9.5% for the moment of inertia around theXandYaxes. In addition, the experimental results for the UAS-S4 inertial values validate the numerical calculation through DATCOM method with a relative error of 6.52% on average.

This paper aims to convert the aerodynamic shape of a conventional aerial subsonic flying body into a glide one by providing a range extension kit and fins. It describes the selection of configuration and airfoils depending on the tactical requirements and flight regimes. The wing and fins sizing is obtained using three different methods subjected to geometric constraints. The first method is an iterative optimization method using linear aerodynamic data. The second method is a genetic algorithm multi-objective function aims to maximize stability, controllability and lift-drag ratio within certain weights using linear aerodynamic data. The third method is a genetic algorithm optimization function integrated with MISSILE DATCOM aims to maximize lift-drag ratio. Then perform a direct uncontrolled six degree of freedom simulation for the three optimized designs and the conventional flying body. Comparing the results of ranges for these bodies reveals that the third method has the best range over the other designs including the conventional flying body.

Stability and manoeuvrability have been the major trade-off in the design of aircraft both for civil or military usage. Hence, the need to design a suitable system for the control of the flight control surfaces (Elevator, Rudder and Aileron) on the aircraft. The flight control system (FCS) design helps resolve both the flying and handling qualities of the Unmanned Aerial Vehicles (UAV); it makes use of Stability Augmentation System (SAS) as a feedback mechanism tool for the FCS. Therefore, this paper discusses the validation of FCS for a Medium Altitude Long Endurance (MALE) UAV, whose service ceiling is 15,000 ft., having an endurance of up to 20 hours. A six degrees of freedom (DOF) aircraft model was used to test both the longitudinal and lateral/directional stabilities of the aircraft. Conceptual design parameters were used to generate both the dimensionless and derivatives of the MALE UAV using the AVL software. The results generated from AVL were validated using the digital DATCOM applications and both applications gave satisfactory results though with slight variations in the values of the coefficients obtained from both applications.

The prime objective of this work is to estimate stability and control derivatives of an airship. The complete, nonlinear mathematical model of aerial vehicles has its aero model as a nonlinear function of angle of attack. This along with the necessity for an exhaustive dataset complicates the estimation procedure. In this work, data are generated by simulating the mathematical model of airship for different trim conditions obtained from continuation analysis. A modular neural network is then trained using back-propagation and Adam optimization algorithm for each aerodynamic coefficient separately. The estimated nonlinear airship parameters are found to be consistent with the DATCOM parameters which were used for open-loop simulation in data generation phase. This validates the proposed methodology and could be extended to estimate airship parameters from real flight data.

— The aim of this work is to model and simulate a Flying Wing Aircraft and compare acquired flight data from two different simulations. The framework proposed constitutes of the development of a graphical model as well as the use of a mathematical model of the aircraft. In order to implement the Software-in-the-Loop (SIL), we used a commercial flight simulation environment, X-plane 10, which simulates the dynamics of the aircraft through the Blade Element Theory, and the software MATLAB/Simulink, which simulates and implement the control laws. We collected the flight data from the simulator, as well as the responses of the mathematical model based on equations of motion and the aerodynamic derivatives, which were computed via the Digital DATCOM software. Both analytical simulation and the SIL simulation were performed with the same flight conditions, validating the longitudinal dynamic of the computational model and allowing further studies of this type of aircraft in future projects.

In this paper, we present the procedure of estimating the aerodynamic coefficients for a commercial aviation aircraft from geometric parameters at low-cruise-flight conditions using US DATCOM (United States Data Compendium) and XFLR software. The purpose of this research was to compare the stability parameters from both pieces of software to determine the efficacy of software solution for a wide-body aircraft at the stated flight conditions. During the initial phase of this project, the geometric parameters were acquired from established literature. In the next phase, stability and control coefficients of the aircraft were estimated using both pieces of software in parallel. Results obtained from both pieces of software were compared for any differences and the both pieces of software were validated with analytical correlations as presented in literature. The plots of various parameters with variations of the angle of attack or control surface deflection have also been obtained and presented. The differences between the software solutions and the analytical results can be associated with approximations of techniques used in software (the vortex lattice method is the background theory used in both DATCOM and XFLR). Additionally, from the results, it can be concluded that XFLR is more reliable than DATCOM for longitudinal, directional, and lateral stability/control coefficients. Analyses of a Boeing 747-200 (a wide-body commercial airliner) in DATCOM and XFLR for complete stability/control analysis including all modes in the longitudinal and lateral directions have been presented. DATCOM already has a sample analysis of a previous version of the Boeing 737, however, the Boeing 747-200 is much larger than the former, and complete analysis was, therefore, felt necessary to study its aerodynamics characteristics. Furthermore, in this research, it was concluded that XFLR is more reliable for various categories of aircraft alike in terms of general stability and control coefficients, and hence many aircraft can be dependably modeled and analyzed in this software.

It is well-known that air transportation of passengers and goods is a large market in the world of aviation. Hybrid buoyant aerial vehicles have the potential to be candidates for such applications. This paper addresses the preliminary static stability analysis for hybrid buoyant aerial vehicles in which a novel concept of ballonet arrangement is applied. A study was performed to assess the suitability of commercial off-the-shelf software for this task. A generic geometric model was created to obtain the lift and drag characteristics of the aircraft and provide geometric parameters. These parameters were input into DATCOM, which was used to model the aerodynamic characteristics and to determine the stability and control derivatives for static analysis and airship performance enhancement. This study demonstrates that the aerodynamic and stability trends for low-speed airships can be obtained and analyzed using DATCOM.

This paper investigates the influence of aerodynamic characteristics for a flying-wing Unmanned Air Vehicle (UAV) subject to asymmetric wing damage. Firstly, the conceptual design of a flying-wing UAV is introduced. The UAV model with right wing damage is constructed, from which the vehicle's physical characteristics, such as mass, center of gravity (C.G.) and moment of inertia are calculated. Next, the aerodynamic model of damaged UAV is established in DATCOM, and the aerodynamic coefficients of UAV under different damage conditions are computed. Based on the obtained aerodynamic coefficients, the effect of asymmetric wing damage for the UAV is analyzed. Furthermore, the influence of the static stability is analyzed according to the aerodynamic derivatives.

Autopilot is an augmented flight control system that is used to reduce the workload of the pilot during long endurance missions. Development of autopilots for Unmanned Aerial Vehicles (UAVs) also needs such a control system before building it for pilotless operation. Presently, the autopilot for fixed wing aircraft is dominated by classical PID controller. However, state space methods provide systematic and more robust control system design. Hence, state space method is considered for an UAV, named as Casper, designed and ground tested at authors University. This paper presents modeling and simulation aspects of UAV and development of pitch control subsystem of autopilot with state space methods. Simulink environment is used with inbuilt aerospace block-sets for flight dynamics and Digital DATCOM to obtain aerodynamic characteristics of UAV. FlightGear simulator is interfaced with Simulink model for visualization of UAV response. Linear Quadratic Regulator (LQR) approach is used in state space method to determine state feedback gain matrix. Comparison of state feedback and PID for pitch control autopilot shows, state feedback controller has better “command following performance”, as it has lesser over shoot of 75%, lower settling time by 0.5 s and 55% lesser controller effort. Response to gust loads of state feedback and PID autopilot are comparatively same. However, transient performance of PID controller is smoother and settling time is 75 s lesser than that is achieved by state feedback controller.

This paper presents the procedures used for estimating the stability and control derivatives of a general aviation canard aircraft from flight data. The maximum likelihood estimation method which accounts for both process and measurement noise was used for the flight data analysis of a four seat canard aircraft, the Firefly. Without relying on the parameter estimation method, several aerodynamic derivatives were obtained by analyzing the steady state flight data. A wind tunnel test, a flight test of a 1/4 scaled remotely controlled model aircraft, and the prediction of aerodynamic coefficients using the USAF Stability and Control Digital Data Compendium (DATCOM), Advanced Aircraft Analysis (AAA), and Computer Fluid Dynamics (CFD) were performed during the development phase of the Firefly and the results were compared with flight determined derivatives of a full scaled flight prototype. A correlation between the results from each method could be used for the design of the canard aircraft as well as for building the aerodynamic database.

Recently, guided missiles applies folding wings to save space. During wing deployment, aero force acting on wing effects significantly on deployment performance, usually aerodynamic coefficient are calculated by CFD analysis. However, Missile Datcom can calculates estimated aerodynamic coefficient very quickly by assuming wing deployment motions as dihedral angle of wing. If missile has external store, wings may need to be folded on top of each other. In this case, one of wing help or interrupt other wing deployment, locking effect. In this study, both effects were included on wing deployment performance analysis to criteria for wings locked condition and formulated wing deploy performance, and compared with wind tunnel test data. Analysis predicted vulnerable wind direction of wing deployment very well.

The main goal of aircraft bombing is its success in the destruction of the required target depending on the correct aiming and the correct moment of release. Different conditions, such as air speed, altitude, angle of release and the presence of wind, have a significant effect on the choice of the release moment to ensure target destruction. In this paper, a Six-degree of freedom (6DOF) mathematical model for the free fall bomb is presented describing the motion of the bomb in the three-dimensional space in the presence of stochastic winds. The aerodynamic coefficients for the free fall bomb are calculated through Missile Datcom computer program, however, these coefficients can be calculated either experimentally using the wind tunnel or a flight test; or mathematically through the Computational Fluid Dynamics (CFD). The stochastic wind affecting the trajectory of the bomb is defined by a velocity spectral represented by the Continuous Dryden Spectral. The contribution in this paper is the modeling and evaluation of the free fall bomb flight simulation in the presence of stochastic wind affecting the bomb flight trajectory.

This paper explores the design, simulation and analysis of three novel Linear Quadratic Gaussian (LQG) control system for a Longitudinal dynamic of a fixed wing mini Unmanned Aerial Vehicle (UAV). Modelling results for the small UAV using Aircraft DIGITAL DATCOM ® are presented. The novelty of the design is from the stand point of the Kalman Filter, with Kalman gain obtained from the solution to a Differential Riccati Equation (DRE) rather than the popular Algebraic Riccati Equation (ARE). The formulated DRE to the Kalman filter design is solved as an Initial Value Problem (IVP) in the MATLAB/Simulink using explicit algorithm ode45. The algorithm converges to a solution of interest with 6671steps. Each step has a covariance matrix hence a different Kalman gain value as the solution tries to converge, after 10 seconds of simulation. Three of 6671 step values are selected based on the observed trajectory of the Kalman gain matrix and the time for the set-point tracking control of the pitch angle to reach a steady-state value. Three of these Kalman Original Research Article gains obtained were used in the design of the linear Kalman filter, which serves as the observer for the Linear Quadratic Regulator (LQR between the three improved LQG characteristics and robustness. Using the robustness properties of the Kalman filter as a benchmark, simulation result shows that all the three improved LQG.

Purpose – This paper aims to clarify the flight dynamics characteristics and improve the flight performance for large-scale morphing aircrafts. With specific focus on the effects of morphing on mass distribution, aerodynamics and flight stability, the study aims to develop the dynamic model, outline the morphing strategies design and evaluate the flight stability in transient stage of morphing. Design/methodology/approach – The mode of relaxing the rigidity condition was opted, which introduced the functions of position of center of mass and moments of inertia with respect to the morphing parameters, and yielded a parameter-dependent flight dynamics model. The morphing strategies were designed by optimizing the morphing parameters with the corresponding performance metric of each mission segment, where the aerodynamics was estimated concurrently by DATCOM. Based on the decoupled and linearized longitudinal parameter-dependent model, the flight stability in transient stage of morphing was evaluated based o...

This paper aimed to predict aerodynamic & stability data for the Safat01 aircraft. Because of its economic and time efficiency, the Digital DATCOM program has been used to predict the stability and control derivatives for the subsonic, low angle of attack (less than 15 degrees) flight regimes. The DATCOM aerodynamic results were very close when compared to CFD. The effect of the nominal centre of the gravity position was considered. According to the computed stability results, the aircraft has been found statically stable and possessing longitudinal dynamic stability. This agrees with the flight tests results conducted on the aircraft which showed stable behaviour.

Digital flight is the study commonly called flight simulation. In flight simulation analysis, the biggest challenges are about preparing the complete databases that can be invoked as result files in future. The ability to simulate in a computer regarding the flight motion must be convinced the governing equations of motion for six-degree of freedom. The accuracy of a flight simulation is depends to a large extent on the quality of an aerodynamic model database prediction. Key objective is to fly and simulate an aircraft in a computer environment with sufficient level of autonomy under the interest variables. In order to realize the goals, a surrogate model is used as the prediction function based on sampling data obtained by the design of experiments. The sample data for models are acquired from the USAF Stability and Control DATCOM which are considered as an exact function, and a database is constructed for two main variables angle of attack and Mach number along the X-force axis, Z-force axis, and pitching moment respectively. Furthermore, a comparison value of exact and predicted function was compared through the error production. The validation for the digital flight was qualified through the mode motion analysis and flying qualities scale to prove the mission level of the flight status.

The goal of this paper is to present a development of a low cost flight simulator, that allows to simulate flight controls failures. Cessna 172 has been chosen as an example of a general aviation aircraft and the flight model has been implemented in Simulink. The model allows for easy integration of an experimental autopilot, using various strategies. Aerodynamic coefficients have been calculated using software called DATCOM. Such approach reduces greatly the cost of development. An approximate inertia matrix has been calculated for proper rigid body dynamics, and a model of engine and propeller is based on actual data from producer. In the end evaluation has been performed, indicating that modeled airplane behaves in a proper way.

A study of the aerodynamic characteristics of two missile configurations has been performed using the data of an experiment, of a semi-empirical aerodynamic prediction code (Missile Datcom) and a RANS flow solver (TAU). The aerodynamic coefficients for different angles of attack and roll angles for Mach numbers of M= 0.85 and 1.2 are compared. Considering the design of the configurations one is rather unconventional with large aspect ratio wings. Whereas the other one has conventional wings with a small aspect ratio. The latter sustains large lift forces at high angles of attack due to the occurrence of vortices while the former performs better at low angles of attack. The comparison of the experimental data with the TAU predicted aerodynamic coefficients shows a good agreement for both configurations in the entire AoA-range. Regarding Missile Datcom, for the unconventional design the accuracy of the predicted data decreases with increasing angle of attack but remains almost constant for the conventional.

In recent years, military and civil aviation have put forward higher and higher requirements to the performance of the aircraft. The morphing aircraft can adapt to a variety of environments, perform a variety of tasks and meanwhile maintain excellent flight performance. This paper researches the dynamic modeling method, aerodynamic parameter analysis and the dynamic characteristic response of morphing aircraft. The physical model of a variable-span variable-sweep morphing aircraft was simplified so as to use Kane Equation modeling. In this paper, the software Datcom is used to analyze the aerodynamic parameters of the morphing aircraft. Then the aircraft's longitudinal dynamic responses during the morphing process are simulated and analyzed. The simulation results show that the speed, height and pitch angle of the morphing aircraft will be greatly changed.

The paper presents the study of unguided rolling projectiles at varying Mach numbers using aerodynamic coefficients. The aerodynamic coefficients are estimated using an aerodynamic prediction code, Missile DATCOM. The predicted dynamic derivatives will determine the design criteria, and also their effect on the design aspects (stability and accuracy) of the projectile. To satisfy the condition of stability for the trajectory of projectile, a model of 6 DOF equations has been used. The result parameters, such as static margin of the projectile, force and moment derivatives in all the three modes, spin rate, angle of attack, and time rate are shown. The accuracy of the rolling moment derivatives is uncertain due to the limitation of DATCOM, but it is the general method applicable at the initial design stage of any projectile and therefore falls in the realm of aerodynamic database. The results address the problems of static and dynamic stability by giving initial perturbations, effect of spin on thrust misalignment, and effect of variation in the geometric parameters during the power on and off cases. Therefore, there is a need for full aerodynamic characterization of the projectile, in which sufficient derivatives are computed for accurate results.

Unmanned Aerial Vehicle modelling has found diverse utilities in both civil and military applications. In order to develop an accurate model of their flight dynamics, it is important to properly estimate their aerodynamics coefficients. For this purpose, several methods are usually applied. This paper presents a methodology to obtain the flight dynamics of an Unmanned Aerial Vehicle, for which its aerodynamics coefficients were found based on its geometrical properties. This methodology was applied to the UAS-S4, designed and manufactured by Hydra Technologies, using DATCOM and TORNADO codes. The aerodynamic model thus found was compared with another model obtained by use of ANSYS Fluent software. The model was completed with a propulsion system developed by use of Javaprop. Results have shown that the obtained model is capable of estimating with accuracy the aerodynamic behaviour of the UAS-S4.

Use of low fidelity tools in designing subscale generic wind tunnel models is usually required to get first-hand knowledge about general trends of aerodynamic and stability parameters. Most of such tools are limited to well-defined conventional aircraft configurations. In the present work, aerodynamic and stability characteristics of a winged hybrid airship is explored at low subsonic speed by using Aircraft Digital DATCOM, which is based on semi-empirical methods for preliminary aircraft geometries. Spheroidal ellipsoidal shaped hull of the airship is modeled in DATCOM along with the geometrical details of wing and empennages. The prediction of zero lift drag coefficient, coefficient of lift and pitching moment is the focus of this paper. Except the drag coefficients, trends of analytical results compare well with experimental data.

The mathematical dynamics model of the tilt-rotor unmanned aerial vehicle (UAV) that has nacelle-fixed auxiliary wings (NFAWs) is presented based on computational fluid dynamics analysis using FLUENT and DATCOM. The advantage of the aerodynamic performance of the NFAW is compared to the performance of the original tilt-rotor UAV in a trim analysis as well as simulation. The inner loop and outer loop of the neural network controller are designed for the tilt-rotor and its NFAW variant. In order to improve the control performance of outer loop, pseudo-control hedging (PCH) is applied to the outer loop as well as the inner loop neural network control. The dynamic inversion on a linear model of the original tilt-rotor at hover conditions is used as a baseline. The sigma-pi neural network (SPNN) adaptation minimizes the error of the inversion model. This error typically occurs due to the use of an approximate tilt-rotor model for helicopter mode instead of the NFAW model throughout the flight envelope from helicopter to airplane mode. The waypoint navigation and the automatic hover guidance are applied to the most outer loop of the neural network controller for the autonomous flight, which consists of nacelle conversion and reconversion as well as automatic take-off and landing. The fast dynamic reference commands generated by the autonomous waypoint guidance are inputted to the outer loop control in order to make the PCH of the outer loop effective. Lastly, the nonlinear simulation results are compared under turbulent wind conditions, in which the NFAW is more negatively affected than the original tilt-rotor model.

Aerodynamic analysis and calculation are very important parts in the aircraft design, and aerodynamic engineering prediction is widely used in the aircraft preliminary design stage. However, traditional aerodynamic engineering prediction causes heavy computation and is time-consuming. The developed software such as DATCOM has the disadvantages of complicated operation and black box structure. To overcome the disadvantages above, we develop the software for aerodynamic engineering prediction based on the aerodynamic characteristics and prediction for aircrafts. There are three parts in this software which are database, calculation module and user interface. The software is verified by a numerical example of one aircraft, and compares with the data of Computational Fluid Dynamics (CFD) and the wind tunnel test. The results show that the calculated results of the aerodynamic engineering prediction and CFD are basically consistent, and the software is able to meet the accuracy demand in the preliminary design phase of the aircraft.

The integrated missile design optimization process is proposed by implementing the aerodynamics database (Aero DB) and tactical missile design (TMD) spreadsheet to obtain a quick and relatively accurate optimal air intercept missile configuration at the conceptual design stage. The Aero DB is constructed to replace an existing aerodynamics analysis module in the TMD spreadsheet and to provide stability and control coefficients as constraints for improving missile range performance based on the body-wing-tail configuration baseline. Sensitivity analysis is performed on an entire missile geometry and flight condition variables to eliminate the small effects of design variables on missile range and constraints under a PHX ModerCenter 10.1 integration environment. The optimal missile configuration shows 27.8% improvement in total range compared with a body-wing-tail configuration baseline while all constraints are satisfied. The proposed integration of the missile design program using Aero DB demonstrates more accurate and reliable results which are validated by high-fidelity analysis ANSYS Fluent 13 on the optimal missile configuration compared with TMD aerodynamics analysis results. The maximum difference between ANSYS Fluent and Missile DATCOM is 11.76% at 10 degrees of AoA compared with 37.97% for TMD aerodynamics analysis and ANSYS Fluent difference.

System identification is an established technique for obtaining a reliable mathematical model of an Unmanned Aerial Vehicle (UAV) using aircraft input output data. In this research, a unique three step procedure is presented for system identification of small scale fixed wing UAV. Flight experiment was conducted with specifically designed maneuvers to identify lateral-directional dynamic model of the aircraft. Initial reference model was developed using United States Data Compendium (DATCOM) and grey box identification was performed in Matlab® system identification toolbox. Model parameters were estimated using Prediction Error Method (PEM) and the estimated parameters were improved iteratively through Adaptive Gauss Newton (GNA) optimization. Model validation was performed and the UAV's aerodynamic coefficients were determined of the UAV. Excellent validation results show the usefulness of the three step identification scheme. The identified lateral model can be practically used for various applications including heading control and autonomous navigation of the UAV.

Missile Datcom (hereinafter referred to as Datcom) is an aerodynamic engineering calculation software developed by the US Air Force Research Laboratory. It is commonly used in the preliminary design of the shape scheme of missiles, aircraft and other aircraft. In this paper, Datcom is used to calculate the aerodynamic parameters of a two-dimensional ballistic correction projectile(D2BCP), and the calculated data is compared with the wind tunnel test data. The analysis results show that Datcom calculates the axial force coefficient (CA) of the projectile with large system deviation, however, the deviation can be eliminated by a certain compensation, thereby obtaining a more accurate CA of the projectile. Datcom calculates the normal force coefficient (CN) with high Mach number accuracy, but there is an unacceptable deviation of the projectile CN calculated at small angle of attack ( α ). The calculation of pitching moment coefficient with α (CMa) is higher, but the error is larger under individual Mach numbers.

Aircraft dynamics and control (ADC) is a core course requirement for an undergraduate program in aeronautical engineering. This article presents an undertaking that reconfigures the instruction of fundamental concepts in ADC by replacing traditional problem-solving approach with a method that relies on a simulation framework comprising of Matlab, USAF DATCOM, and flight simulation software. The typical challenges experienced by students including their inability to visualize complicated, multi-modal aircraft were overcome by the new method thus enhancing the student's learning experience. The implemented approach improved students' motivational belief and its calibration, their use of active learning strategies and actual performance. © 2013 Wiley Periodicals, Inc. Comput Appl Eng Educ 23:63–71, 2015; View this article online at wileyonlinelibrary.com/journal/cae; DOI 10.1002/cae.21579

In this paper, preliminary effort of pre-flight test lateral control method study is presented. Parameters and coefficients of candidate UAV (Unmanned Aviation Vehicle) are calculated for modeling and linearization. For controller design, dynamic inversion is used to cancel the dynamic coupling between attitudes. Stability of inversion loop controller which applied on non-ideal inversion is studied. Finally, simulation is present to compare the PI and nonlinear PI controller based on dynamic inversion.

Because the missile is autonomous, its control system must provide adequate flight stability while ensuring sufficient response to track commands. This paper is devoted to investigate the usefulness of the classical and modern control techniques for autopilot design for different evaluation approaches. The present work is concerned with improving theperformance of a surface-to-surface controlled aerodynamically guided missile system via both classical PID and predictive autopilots. The design and analysis necessitate somehow accurate system model with different uncertainties via 6-DOF simulation. The governing differential equations of the missile motion are derived with the aerodynamic model of themissile constructed by means of the Missile Datcom software. After obtaining the required aerodynamic stability derivatives using the generated aerodynamic data, the necessary transfer functions are determined based on the equations of the missile motion. Next, the normal acceleration autopilot is designed using the determined transfer functions. The autopilot is designed to realize the command signals generated by the guidance laws which are in theform of normal acceleration components. Using the entire system model, the computer simulations are carried out using the Matlab-Simulink software where the classical and predictive autopilots are compared via time response along the flight path

This paper presents an aerodynamic study of “ × − + ” finned configurations for theater ballistic missile target (TBMT) which plays a key role to evaluate the Missile Defence System. The semi-empirical method (MISSILE DATCOM) and CFD method were selected to calculate the aerodynamic coefficients and flowfield for multiple sub-models. A wind tunnel experiment was conducted to explore the aerodynamic characteristics of this TBMT model in supersonic condition. Both computational methods are capable of predicting the aerodynamic coefficients of TBMT. The experimental results also showed that the “all stages” modelʼs pressure center only had small changes when α is increasing and the TBMTʼs maneuverability would increase after the first separation. The flowfield visualizations showed that the front finʼs trailing vortices had no interference with tail fins in small angle of attack condition. The “ × − + ” finned aerodynamic configuration can meet the need for high stability to lower the cost of the guidance and control system in the TBMT.

This paper designs a morphing UAV with variable span and variable sweep based on BQM-34 “Firebee” as a variant platform. By using a software called Datcom to calculate the aerodynamic coefficients between different variable configurations and considering the flow field hysteresis effect during the variant motion process, the aerodynamic parameter model of the morphing aircraft is established. The morphing aircraft can be simplified as a multi-rigid body system, and then Kane method is adopted to establish dynamic model of it. The longitudinal motion simulation of the model indicates that the model built in this paper is corresponding to the fixed wing aircraft dynamics response as well as it can reflect the dynamic characteristics of the variant process. Eventually ,it provides a precise model for control system of a morphing UAV.