Dynamic Stability Derivatives Research Articles

Numerical Evaluation of Aircraft Aerodynamic Static and Dynamic Stability Derivatives by a Mid-Fidelity Approach

The present study aimed to investigate the capability of mid-fidelity aerodynamic solvers in performing a preliminary evaluation of the static and dynamic stability derivatives of aircraft configurations in their design phase. In this work, the mid-fidelity aerodynamic solver DUST, which is based on the novel vortex particle method (VPM), was used to perform simulations of the static and dynamic motion conditions of the Stability And Control CONfiguration (SACCON): an unmanned combat aerial vehicle geometry developed by NATO’s Research and Technology Organisation (RTO), which is used as a benchmark test case in the literature for the evaluation of aircraft stability derivatives. Two different methods were exploited to extract the dynamic stability derivative values from the aerodynamic coefficient time histories that were calculated with DUST. The results for the mid-fidelity approach were in good agreement with the obtained experimental data, as well as with the results obtained using more demanding high-fidelity CFD simulations. This demonstrates its suitability when implemented in DUST for predicting the static and dynamic behavior of airloads in different conditions, as well as in reliably predicting the values of stability derivatives, with the advantage of requiring limited computational effort with respect to classical high-fidelity numerical approaches and the use of wind tunnel tests.

High-Efficiency Simulation of Dynamic Stability Derivatives Based on a Particle Swarm Optimization and Long Short-Term Memory Network (PSO-LSTM) Coupling Aerodynamic Model

A new high-efficiency method based on a particle swarm optimization and long short-term memory network is proposed in this study to predict the aerodynamic forces in an unsteady state. Based on the predicted aerodynamic forces, the dynamic derivative is further calculated. Using particle swarm optimization to optimize the hyper-parameters of a neural network, the long short-term memory network prediction model can be constructed according to the known simulating aerodynamic data to predict the aerodynamic performance of aircraft in unknown states. By coupling the least-squares method in an aerodynamic derivative model, the dynamic derivative can be quickly obtained. Unsteady motion of NACA 0012 airfoil was taken as the research example to verify this method, and its longitudinal combined dynamic derivatives were predicted and compared with CFD simulation results. The results show that the dynamic derivatives predicted by the PSO-LSTM method have high accuracy, with an error of no more than 1% compared to CFD, and a 70% improvement in efficiency. The method proposed in this study has good generalization ability and can realize fast and accurate prediction of dynamic derivatives with a small number of samples.

Small-Perturbation Analysis of Airplane Dynamics With Dynamic Stability Derivatives Redefined

The first successful powered and controlled heavier-than-air flight and the subsequent meteoric rise of aviation in the 20th century also brought with it mathematical models to understand the phenomenon of flight. Though successful to a large extent as is obvious from their longevity, these models still suffer from the odd anomalous result. Upon careful examination, the source of this problem can be traced to an inappropriate choice of dynamic stability derivatives that is inconsistent with the aerodynamics. In order to correctly distinguish between the different mechanisms that are responsible for the aerodynamic forces and moments generated due to body angular rates in flight, this paper refines the existing model for flight mechanics by redefining the dynamic stability derivatives with respect to, i) relative rate of rotation between the body and wind axes, and ii) the rate of rotation of the wind axis with respect to the earth axis, sometimes called flow curvature effect. Further, new literal approximations to the longitudinal and lateral-directional modes have been obtained in terms of the redefined dynamic stability parameters. The new approximations, for instance, correctly relate the short-period natural frequency to the longitudinal static stability parameter Cmα.

Nonlinear system control analysis and optimization using advanced Pigeon-Inspired optimization algorithm

The Pigeon-Inspired optimization (PIO) algorithm is a novel intelligent optimization algorithm inspired by birds’ behavior as their travel. This; behavior modeled to be used for solving many optimization problems in different fields. However; it always suffers from unstable behavior when used with nonlinear; time-varying systems. In; this paper, this algorithm is adapted to calculate the optimum controller gains for roll and pitch channels in a guided tactical missile. The; vehicle model is presented in a nonlinear; form and then shown in a linearized form for the sake of an autopilot design. The PIO; algorithm is supported and accompanied by an adaptive algorithm to determine the initial states and constraints for the PIO algorithm to enhance the behavior of the optimization algorithm to speed up the convergence rate to reach an optimum and feasible solution. Also; an estimation function is incorporated to estimate model parameters variation such as dynamic pressure, stability derivatives, and mass properties. Meanwhile; a comparative analysis is carried out with original PIO and particle swarm optimization algorithms, utilizing a non-linear; model with the presence of noise source and disturbance to ensure the ability of the algorithm to make the autopilot robust and stable against several sources of uncertainties.

Stability Derivatives of Various Lighter-than-Air Vehicles: A CFD-Based Comparative Study

An aerostat with a single tether is proposed for the application of wind measurements at low altitudes. In the current study, the aerodynamic model parameters (stability derivatives) of the aerostat are investigated based on a CFD-based approach. The static, as well as the dynamic stability derivatives of the aerostats are presented. The calculation of the dynamic stability derivatives involves the simulation of the oscillations of the aerostats in their axial direction (surge), the vertical direction (heave) and angular motions with respect to the lateral direction (pitch). A forced sinusoidal oscillation is used for the simulation of the aerostat, and one stable period of oscillation is taken for the derivatives’ extraction. Four different aerostats are considered for the current study with four different angles of attack. The Zhiyuan aerostat, HAA aerostat, NPL aerostat and GNVR aerostat are the aerostats considered for this study. The stability derivative results obtained for the four aerostats are analyzed and compared with respect to their geometrical features. From the static aerodynamic characteristics, the Zhiyuan aerostat shows better performance than the other aerostats in terms of the lift–drag ratio. The dynamic stability derivatives of the Zhiyuan aerostat suggest its application as the proposed low-altitude wind measurement system.

Estimation and Separation of Longitudinal Dynamic Stability Derivatives with Forced Oscillation Method Using Computational Fluid Dynamics

This paper focuses on estimating dynamic stability derivatives using a computational fluid dynamics (CFD)-based force oscillation method, and on separating the coupled dynamic derivatives terms obtained from the method. A transient RANS solver is used to calculate the time history of aerodynamic moments for a test model oscillating about the center of gravity, from which the coupled dynamic derivatives are estimated. The separation of the coupled derivatives term is carried out by simulating simple harmonic oscillation motions such as plunging motion and flapping motion which can isolate the pitching moment due to AOA rate (Cmα˙) and the pitching moment due to pitch rate (Cmq), respectively. The periodic motions are implemented using a CFD dynamic mesh technique with user-defined function (UDF). For the validation test, steady and unsteady simulations are performed on the Army-Navy Finner Missile model. The static aerodynamic moments and pressure distribution, as well as the coupled dynamic derivative results from the pitching oscillation mode, show good agreement with the previously published wind tunnel tests and CFD analysis data. In order to separate the coupled derivative terms, two additional harmonic oscillation modes of plunging and flapping motions are tested with the angle of attack variations from 0 to 85 degrees at a supersonic speed to provide real insight on the missile maneuverability. The cross-validation study between the three oscillation modes indicates the summation of the individual plunging and flapping results becoming nearly identical to the coupled derivative results from the pitching motion, which implies the entire set of coupled and separated dynamic derivative terms can be effectively estimated with only two out of three modes. The advantages and disadvantages of each method are discussed to determine the efficient approach of estimating the dynamic stability derivatives using the forced oscillation method.

Self-Induced Roll–Yaw Oscillations of an Aircraft Model in a Wind Tunnel

Wind tunnel oscillatory testing is a technique conventionally used for the measurement of dynamic aerodynamic loads on subscale aircraft models. Oscillatory testing is commonly used for identification of dynamic stability derivatives for flight mechanics models. Model oscillation is normally driven by a rig drive mechanism, or with the use of its control surfaces. This paper presents a new test technique in which self-induced coupled roll and yaw oscillations are driven purely by the aircraft’s natural modes of motion when constrained in translation and in pitch rotation. This method does not require external mechanisms to drive the motion, nor the use of aircraft control surface inputs to sustain the oscillation. Without these sources of error, it is potentially advantageous for identification of dynamic stability derivatives using parameter estimation techniques. The mechanics of and requirements for this self-induced oscillation are discussed. Computational simulation and dynamic wind tunnel tests of a subscale BAe Hawk model are used to demonstrate the self-induced roll–yaw oscillations and their results are compared. Effects of varying parameters such as wind tunnel speed and the addition of spring stiffness onto the aircraft rotation axes on the frequency and stability of the oscillation are investigated.

Low speed lateral-directional dynamic stability analysis of a hypersonic waverider using unsteady Reynolds averaged Navier Stokes forced oscillation simulations

Hypersonic waveriders have the potential to significantly reduce travel times on long haul civilian transport routes. The design of hypersonic aircraft is heavily influenced by the aerodynamic efficiency at the cruise Mach number, resulting in less than ideal geometries for subsonic flight. Waverider aerodynamics and stability in the low speed regime is rarely investigated and not well understood, but is crucial for horizontal take-offs and landings. To date, low speed analyses of waverider shapes have been confined to static investigations, with limited work on longitudinal dynamics. No studies outlining the lateral-directional dynamic stability and aerodynamic derivatives has ever been published. This paper presents results for the low speed propelled variant of the Mach 8 HEXAFLY-INT waverider, which was subjected to forced oscillations in roll, yaw and side-to-side translations. This was achieved using unsteady Reynolds Averaged Navier Stokes simulations. Tests were conducted at a speed of 20 m/s, which correlates to a Reynolds number of approximately 1.5×106. The dynamic motion was at a forced frequency of 1 Hz with angle amplitudes of 1 degree. The vehicle was analysed through an angle of attack range from -5 to 15 degrees in 5 degree increments. Results for the nominal centre of gravity location at 44.4% of the vehicle length show values for the dynamic derivatives within expected ranges, apart from the rolling moment due to yaw rate below approximately 2 degrees angle of attack. A similar finding for the dihedral derivative in static tests concluded that the low mounted wing with anhedral results in instabilities at low angles of attack. Investigations conducted at a centre of gravity location of 53.1% of the vehicle length, the aft static stability limit, also showed data within normal ranges. For these cases, the derivative magnitudes were lower, which indicates decreased damping compared to the nominal centre of gravity location. Some significant impact was seen on the roll damping derivative due to lowering the centre of gravity. This is from a combination of non-linear wing side force variance with angle of attack and the reduction of the moment arm, which resulted in less damping. The derivatives generally showed little sensitivity to changes in the oscillation frequency, with frequency rates ranging from 0.5 to 2 Hz tested. Overall, the results from this study highlighted the complex combinations of rate induced forces which contribute to the dynamic behaviour of the aircraft.

Evaluation of Dynamic Stability Derivatives for Hypersonic Vehicle

The dynamic stability derivative is the key parameter of the aircraft control system and the boundary analysis of the dynamic stability criterion. The accurate evaluation of dynamic derivative is of great importance to the design and flight of the aircraft. A method to predict the derivative of pitch damping is presented. The unsteady aerodynamic force is obtained by solving the unsteady flow field, and the static and dynamic derivatives of the aircraft are obtained by using integral method and least square method. By comparing the aerodynamic derivatives at different equilibrium positions and different centers of gravity, it can be seen that the calculated results are in good agreement with the experiment. It indicates that the present method has satisfactory order and efficiency and be suitable for engineering problems.

Simulation on the dynamic stability derivatives of battle-structure-damaged aircrafts

Accurately evaluating the aerodynamic performance of a battle-structure-damaged aircraft is essential to enable the pilot to optimize the flight control strategy. Based on CFD and rigid dynamic mesh techniques, a numerical method is developed to calculate the longitudinal and longitudinal-lateral coupling forces and moments with small amplitude sinusoidal pitch oscillation, and the corresponding dynamic derivatives of two fragment-structure-damaged and two continuous-rod-damaged models modified from the SACCON UAV. The results indicate that, at the reference point set in this paper, additional positive damping is generated in fragment-damaged configurations; thus, the absolute values of the negative pitch dynamic derivative increase. The missing wingtip induces negative pitch damping on the aircraft and decreases the value of the pitch dynamic derivative. The missing middle wing causes a noticeable increase in the absolute value of the pitch dynamic derivative; the missing parts on the right wing cause the aircraft to roll to the right side in the dynamic process, and the pitch-roll coupling cross dynamic derivatives are positive. Moreover, the values of these derivatives increase as the damaged area on the right wing increases, and an optimal case with the smallest cross dynamic derivative can be found to help improve the survivability of damaged aircraft.

Review of Numerical Simulations on Aircraft Dynamic Stability Derivatives

The dynamic stability derivatives of flight vehicle are directly related to unsteady aerodynamics during maneuverability, and are considered as key parameters to both the aerodynamics, control system design and flight qualities evaluation. As the rapid development of new conceptual aircraft configurations, higher precision and efficiency calculations of the dynamic stability derivatives are urgently required. Limited by high costs and risks, the traditional experimental ways of flight test and wind tunnel test cannot be widely used, thus the numerical calculation on dynamic stability derivatives have become the primarily approaches. This paper reviews the numerically methods applied in the estimation of aircraft dynamic stability derivatives, includes the early analytical method, the empirical and semi-empirical methods and the widely used modern time domain and frequency domain methods, deeply highlighting the advantages and drawbacks of these methods of the actual applications on various aircrafts. The early analytical method by using potential theory can be used to calculate dynamic stability derivatives and will never work for current geometries. The empirical and semi-empirical methods are available with simple mathematical model and existed data. They have the advantage to be simple and rapid to compute and can be well adapted for initial evaluation of aircraft conceptual design. The time domain analysis based on solving Euler or NS equations with CFD technique are the most widely used methods to obtain the aircraft dynamic stability derivatives. With a variety of different strategies, they can calculate the combined and single dynamic derivatives to satisfy the demand of aircraft design in each stage. Even they are accurate and better in adaptability, the huge time cost on the periodic unsteady flow limits the application. To overcome this problem, the frequency domain methods based on harmonic oscillation are developed. They only use the results at several sample points during the unsteady cycle to reconstruct the periodic unsteady flows to further efficiently obtain the dynamic stability derivatives. These frequency domain methods are currently available only in harmonic oscillation cases. This paper also discusses and analyses the existing problems and possible development directions of the numerical methods to calculate aircraft dynamic stability derivatives from four aspects: theory, calculating elaboration, efficiency and accuracy, and application.

Control device effectiveness studies of a 53∘ swept flying wing configuration. Experimental, computational, and modeling considerations

The present investigation covers studies of different control devices on several 53∘ swept flying wing configurations without a vertical tail plane. The wings considered share the same planform but differ in their spanwise profile shapes. The planform, developed under the NATO AVT-251 Task Group, is referred to as the MULDICON (or MULti-DIsciplinary CONfiguration). The objectives of this article are twofold: (1) to design yaw control surfaces for the MULDICON wing using an experimental/computational approach and (2) to develop aerodynamic models that rapidly and accurately predict the effectiveness of control surfaces over a wide range of flight conditions. The yaw control surface design (position and size) should provide sufficient yaw moment with almost no contribution in roll and pitch moment. To identify such concepts, a number of preliminary experiments on a generic flying wing configuration have been conducted. Two promising concepts from wind tunnel tests were then numerically examined for being implemented in the MULDICON baseline wing. Concepts with spoilers and a split flap were specifically considered. For medium to high angles of attack, the flow topology of the baseline wing is dominated by a vortical flow field on the upper outer wing. This leads to interactions between vortex and the control device which influences the flow and the attitude of the control device on the upper wing side. Furthermore, this work considers developing aerodynamic models for predicting stability derivatives of several MULDICON designs over a wide range of flight conditions. Aerodynamic loads models are only developed for normal force and pitch moment coefficients, however, the developed approach can easily being extended to include lateral aerodynamic coefficients as well. Quasi-steady models of this work are power series expansions of traditional linear aerodynamic models to capture nonlinear effects. Additionally, the models can estimate static and dynamic stability derivatives and the control surface powers.

Prediction of Aircraft Dynamic Stability Derivatives Using Time-Spectral Computational Fluid Dynamics

Prediction of Aircraft Dynamic Stability Derivatives Using Time-Spectral Computational Fluid Dynamics

Numerical study of saddle-node bifurcation for longitudinal flight with CFD/RBD technique

The square cross section missile was found instability in the longitudinal direction at a certain Mach number in a range of angle of attack both in CFD and wind tunnel test results. To deeply understand the flight behavior and the aerodynamics of the missile, the coupled Computational Fluid Dynamic/Rigid Body Dynamics (CFD/RBD) approach was employed to simulate the free flight process, and the predict technique of dynamic stability derivatives was founded to analyze the types of equilibrium points. The study shows that the type of trim angle of attack is turned from stable node to an unstable saddle with the increase of Mach number, known as the saddle-node bifurcation. There are two extra stable nodes are produced and located on the left and right sides of the saddle point when the Mach number is bigger than the transition Mach number. Further simulation results of free pitching motions reveal the decrease of the attraction region of the two extra stable modes, which makes the pitching motion more complicated and sensitive. A slight change of the initial angle of attack may make the missile move from one trim point to another.

AEROM: NASA's Unsteady Aerodynamic and Aeroelastic Reduced-Order Modeling Software.

The origins, development, implementation, and application of AEROM, NASA's patented reduced-order modeling (ROM) software, are presented. Full computational fluid dynamic (CFD) aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers using the NASA FUN3D CFD code, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. The method and software have been applied successfully to several configurations including the Lockheed-Martin N+2 supersonic configuration and the Royal Institute of Technology (KTH, Sweden) generic wind-tunnel model, among others. The software has been released to various organizations with applications that include CFD-based aeroelastic analyses and the rapid modeling of high-fidelity dynamic stability derivatives. Recent results obtained from the application of the method to the AGARD 445.6 wing will be presented that reveal several interesting insights.

Hypersonic dynamic stability of wave riders

PurposeThe purpose of this paper is to obtain close form expressions for the dynamic stability of conical wave riders with flat surfaces which could be equipped with lifting surfaces on its plain flat surface. Numerical simulation would require very large meshes to resolve flows at subscale level and the experimental evaluations would be equally difficult, requiring expensive measurement facilities with challenging procedures to secure such vehicles in confined test sections to obtain satisfactory wind on and wind off oscillations.Design/methodology/approachThe design method uses appropriate pressure fields using small disturbance theory, which, in turn, is perturbed using the unsteady shock expansion theory to recover suitable expressions for the dynamic stability behaviour.FindingsIt was observed that the dynamic stability of the standard half-cone-type wave riders with flat upper surfaces deteriorates with the axis position measured from the pointed apex reaching a minimum at around x/co = 0.666. The half-cone wave rider with flat upper surfaces is dynamically less stable than a pure cone.Research limitations/implicationsThe method is typically less accurate when the similarity parameter M∞θ ≤ 1 = 1 or if the angle of attack is not small.Practical implicationsWith renewed interest in hypersonic, future hypersonically would be designed as fast lifting bodies whose shapes would be close to the configurations of hypersonic wave riders, especially if they are designed to operate at upper atmosphere altitudes.Originality/valueThe analytic approach outlined in this paper for the evaluation of dynamic and static stability derivatives is original, drawing from the strengths of the small disturbance theory and shock expansion techniques. The method is particularly important, as there are no reported theoretical, numerical or experimental results in the literature.

Computational Modeling of Dynamic Stability Derivatives for Generic Airfoils

This paper presents a method for the computation of the static and dynamic stability derivatives of generic airfoils using high fidelity Computational Fluid Dynamics. Aerodynamic coefficients are calculated for NACA 0012 airfoil and flat plate at different angles of attack. Results of lift coefficient are validated with experimental data. Static and dynamic stability derivatives are calculated by oscillating the airfoil geometry at suitable frequency. Simulations are performed at various flight conditions in terms of angles of attack, frequencies and oscillation amplitudes. The aim of the work is to decipher the behaviour of longitudinal damping derivatives used in flight mechanics through CFD. This approach enables the efficient and accurate computation of dynamic derivatives. Calculations are done for constant air velocity altering only the angle of attack. Inviscid model is tested since its results nearly match with experimental data. The simulations show that the nonlinear characteristics of the stability derivatives are captured by varying angle of attack.

New Systematic CFD Methods to Calculate Static and Single Dynamic Stability Derivatives of Aircraft

Several new systematic methods for high fidelity and reliability calculation of static and single dynamic derivatives are proposed in this paper. Angle of attack step response is used to obtain static derivative directly; then translation acceleration dynamic derivative and rotary dynamic derivative can be calculated by employing the step response motion of rate of the angle of attack and unsteady motion of pitching angular velocity step response, respectively. Longitudinal stability derivative calculations of SACCON UCAV are taken as test cases for validation. Numerical results of all cases achieve good agreement with reference values or experiments data from wind tunnel, which indicate that the proposed methods can be considered as new tools in the process of design and production of advanced aircrafts for their high efficiency and precision.

강제진동 기법을 이용한 무미익 비행체의 동안정 미계수 측정

강제진동 기법을 이용한 무미익 비행체의 동안정 미계수 측정

磁力支持を用いた火星探査カプセルの動安定計測

Almost all re-entry capsules such as HAYABUSA re-entry capsule are dynamically unstable in transonic region. We researched dynamic stability characteristics of Mars entry capsule (70º sphere-cone capsule model) at low speed. In order to avoid significant sting interference, the tests were conducted using magnetic suspension balance system (MSBS). Angle of attack and Reynolds number effects on drag coefficients were measured. Dynamic stability derivatives were measured by forced oscillation method. Drag coefficients were barely affected by Reynolds number and angle of attack. The dynamic stability derivatives of Mars entry capsule were small negative value. These results showed that Mars entry capsule was slightly dynamically unstable.