Aerodynamic Database Research Articles

Aerodynamic modelling of fin stabilised underslung loads

AbstractBox-like slung loads exhibit periodic yaw response instabilities, while carried externally by a helicopter. When coupled with the slung load longitudinal and lateral pendulum motions, these instabilities result in significant pendulum oscillations of the load. High amplitude oscillations lead in many cases to the limiting of a load’s flight envelope. Using wind tunnel and flight tests, rear mounted fins were previously demonstrated as efficient means for stabilisation of a problematic load. However, the lack of a proper analytical model of the stabilised load’s aerodynamic characteristics, led to a trial and error development process, without an appropriate physical understanding of the stabilisation problem. The present paper describes a method for the aerodynamic modeling of fins stabilised slung loads based on a limited number of simple static wind-tunnel tests. The resulting database is incorporated in a dynamical slung load simulation that shows good agreement with dynamic wind-tunnel tests. The applicability of the proposed method is demonstrated, by the calculation of stabilised loads aerodynamic databases for interim fin inclination angles not covered by tests.

The impact of control allocation on trim drag of blended wing body aircraft

The trim drag of conventional aircraft configurations in cruise flight is in the order of 0–1% of the total aircraft drag. Blended wing body aircraft typically have multiple redundant flight control surfaces at the trailing edge with only a limited moment arm with respect to the center of gravity. This aircraft configuration therefore potentially has much more trim drag than conventional aircraft. The redundant set of control surfaces should therefore be deflected in an optimal way to minimize trim drag. A study is conducted to investigate the impact of three control allocation methods on the trim drag of a blended wing body aircraft configuration at relatively low airspeeds. The trim drag is measured by a wind tunnel test campaign. Results indicate that the type of control allocation method has a large influence on the trim drag. This effect must be taken into account in the early design stages of blended wing body aircraft since it will significantly affect the overall aircraft drag. Most control allocation schemes are based on a linear representation of the aerodynamics. Results for the test case indicate that a nonlinear aerodynamic database is required to assess the trim drag with sufficient accuracy. The two main nonlinear effects that must be included are the variations of control derivatives with (1) angle of attack and (2) control deflection. Control surface interaction effects have only a minor influence. The methodology developed to assess trim drag and to select a control allocation algorithm can be directly applied in the early design stages when numerical aerodynamic analysis tools are used instead of wind tunnel models.

Numerical Investigation of Jet Interaction for Missile with Continuous Type Side Jet Thruster

A continuous type side jet controller which has four nozzles with thrust control devices was considered. It is deployed to a missile for high maneuverability and fast controllability in the terminal guidance phase. However, it causes more complex aerodynamic jet interactions between the side jet and the supersonic free stream than does the conventional impulse type side jet with a small single thruster. In this paper, a numerical investigation of the jet interference effects for the missile equipped with a continuous type side jet thruster is presented. A three-dimensional flow field was simulated by using a commercial unstructured-based CFD solver. The numerical simulation method was validated through comparison with wind tunnel test results for the single jet. The method of defining jet direction for this type of side jet control to minimize simulation cases was also introduced. Flow fields investigation and jet interaction effects for various flow conditions, jet pressure ratios and defined jet direction conditions were performed. From the numerical simulation for the continuous type side jet, extensive aerodynamic interference data were obtained to construct an aerodynamic coefficients database for precise missile control.

KF-16의 DOE기반 정적 공력 데이터베이스 구축을 위한 실험적 연구

Wind tunnel testing to construct static aerodynamic database of KF-16 was conducted for preceding research of design of experiments in wind tunnel testing. The test model is KF-16 scaled 1/33 and it has horizontal tail, flaperon, and rudder. The experiments consist of one experiment for analyzing aerodynamic coefficients under whether or not horizontal tail is present and four experiments for analyzing aerodynamic coefficients of changes of deflection angle in control surface which are flap, flaperon, rudder, and horizontal tail. After conducting wind tunnel testing, the experimental results show that the control surface changes have a great effect on Aerodynamic characteristics.

Correlation Analysis of Aerodynamic Forces acting onTall Buildings with Various Side Ratios

본 연구는 고층건축물의 기본설계 단계에서 평면형태를 결정할 때, 응용할 수 있도록 다양한 형상을 가진 고층건축물에 작용하는 풍력 및 풍응답에 대한 데이터베이스를 구축하는데 목적이 있다. 따라서 본 논문에서는 풍력데이터베이스의 초기단계로, 높이 200m이상의 고층건축물의 변장비(=0.33, 0.50, 0.77, 0.83, 0.91, 1.0, 1.1, 1.2 ,1.3 ,2.0, 3.0) 변화에 따른 풍력의 특성을 조사하고 기존 문헌과의 비교·검토를 통해 실험결과를 검증하였다. 또한 풍하중조합의 관점에서 풍방향-풍직각방향, 풍방향-비틀림방향 및 풍직각방향-비틀림방향 간의 풍력의 상호상관에 대해서 검토하였다. 실험결과, 풍방향 및 풍직각방향에 대한 평균 및 변동 모멘트계수와 풍직각방향 및 비틀림방향 변동모멘트계수의 파워스펙트럼밀도의 결과는 기존 연구결과와 잘 부합되는 것으로 나타났다. 또한, 지표면조도의 변화에 따른 풍방향, 풍직각방향 및 비틀림방향에 대한 풍력의 특성은 큰 변화가 없는 것으로 나타났다. 시간영역과 주파수영역에서의 변동모멘트의 상호상관계수 분석결과, 풍직각방향-비틀림방향 변동모멘트계수의 상호상관계수가 높게 나타나는 경향을 보였다.The objective of this study is to construct the TPU aerodynamic database with wind tunnel test data of overall wind loads and responses on tall buildings. In this study, wind tunnel tests were conducted to investigate characteristics of wind forces and the effect of wind load combination by cross-correlation analysis among along-wind overturning moment, across-wind overturning moment and torsional moment on a tall building with various side ratios(=0.33, 0.50, 0.77, 0.83, 0.91, 1.0, 1.1, 1.2 ,1.3 ,2.0 and 3.0) for different terrain roughnesses. The results of wind tunnel tests were compared with those of past literatures. As a result, there was no significant effects of changing of terrain roughnesses on moment coefficients and power spectral densities of across-wind overturning moment coefficients and torsional moment coefficients with various side ratios. Further, these results were good agreement with those of past literatures. From cross-correlation analysis, the across-wind overturning moment coefficients were highly correlated with the torsional moment coefficients. The results of this study will be helpful for practical designers in preliminary design stage.

Surrogate Model of Aerodynamic Model toward Efficient Digital Flight

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.

Prediction of negative peak wind pressures on roofs of low-rise building

In this paper, a probability distribution which is consistent with the observed phenomenon at the roof corner and, also on other portions of the roof, of a low-rise building is proposed. The model is consistent with the choice of probability density function suggested by the statistical thermodynamics of open systems and turbulence modelling in fluid mechanics. After presenting the justification based on physical phenomenon and based on statistical arguments, the fit of alpha-stable distribution for prediction of extreme negative wind pressure coefficients is explored. The predictions are compared with those actually observed during wind tunnel experiments (using wind tunnel experimental data obtained from the aerodynamic database of Tokyo Polytechnic University), and those predicted by using Gumbel minimum and Hermite polynomial model. The predictions are also compared with those estimated using a recently proposed non-parametric model in regions where stability criterion (in skewness-kurtosis space) is satisfied. From the comparisons, it is noted that the proposed model can be used to estimate the extreme peak negative wind pressure coefficients. The model has an advantage that it is consistent with the physical processes proposed in the literature for explaining large fluctuations at the roof corners.

Meta-model based optimization of a large diameter semi-radial conical hub engine cooling fan

Turbomachinery design is an iterative process that can be time-consuming and expensive, especially when an extensive knowledge of the performance envelope is required. The approach described in the present paper can significantly cut the turnaround times down without jeopardizing the accuracy of the final result. A parameterization technique based on radial basis functions (RBF) is used and Reynolds Averaged Navier-Stokes (RANS) simulations are subsequently performed on a set of selected morphed meshes, the goal of which is to produce an aerodynamic database containing first-order, second-order and second-order cross derivatives of objectives with respect to parameters. New solutions, corresponding to any variations of the selected parameters, can thus be extrapolated thanks to the information included in the aforementioned database. In this way, a meta-model is built and can be easily explored by a genetic algorithm. This approach has been experimented on a new concept of engine cooling fan featuring low torque and high efficiency. A reference fan design has been adapted for the particular surrounding of the vehicle underhood, where the downstream flow is radially deviated from its axis by the engine. The optimization process has resulted in an efficiency improvement of three points for one of the obtained optima.

Building Data Fusion Surrogate Models for Spacecraft Aerodynamic Problems with Incomplete Factorial Design of Experiments

This work concerns a construction of surrogate models for a specific aerodynamic data base. This data base is generally available from wind tunnel testing or from CFD aerodynamic simulations and contains aerodynamic coefficients for different flight conditions and configurations (such as Mach number, angle-of-attack, vehicle configuration angle) encountered over different space vehicles mission. The main peculiarity of aerodynamic data base is a specific design of experiment which is a union of grids of low and high fidelity data with considerably different sizes. Universal algorithms can’t approximate accurately such significantly non-uniform data. In this work a fast and accurate algorithm was developed which takes into account different fidelity of the data and special design of experiments.

Unified Framework for Training Point Selection and Error Estimation for Surrogate Models

A unified framework for surrogate model training point selection and error estimation is proposed. Building auxiliary local surrogate models over subdomains of the global surrogate model forms the basis of the proposed framework. A discrepancy function, defined as the absolute difference between response predictions from local and global surrogate models for randomly chosen test candidates, drives the framework, thereby not requiring any additional exact function evaluations. The benefits of this new approach are demonstrated with analytical test functions and the construction of a two-dimensional aerodynamic database. The results show that the proposed training point selection approach improves the convergence monotonicity and produces more accurate surrogate models compared to random and quasi-random training point selection strategies. The introduced root-mean-square discrepancy and maximum absolute discrepancy exhibit close agreement with the actual root-mean-square error and maximum absolute error, respectively, and are therefore proposed as a measure for the approximation accuracy of surrogate models in applications of practical interest. Multivariate interpolation and regression is employed to build local surrogates, whereas kriging and polynomial chaos expansions serve as global surrogate models in demonstrating the applicability of the proposed framework.

Prediction of Aircraft’s Longitudinal Motion Based on Aerodynamic Coefficients and Derivatives by Surrogate Model Approach

The accuracy of a flight simulation is highly dependent on the quality of the aerodynamic database and prediction accuracies of the aerodynamic coefficients and derivatives. A surrogate model is an approximation method that is used to predict unknown functions based on the sampling data obtained by the design of experiments. This model can also be used to predict aerodynamic coefficients/derivatives using several measured points. The objective of this paper is to develop an efficient digital flight simulation by solving the equation of motion to predict the aerodynamics data using a surrogate model. Accordingly, there is a need to construct and investigate aerodynamic databases and compare the accuracy of the surrogate model with the exact solution, and hence solve the equation of motion for the flight simulation analysis. In this study, sample datas for models are acquired from the USAF Stability and Control DATCOM, and a database is constructed for two input variables (the angle of attack and Mach number), along with two derivatives of the X-force axis and three derivatives for the Z-force axis and pitching moment. Furthermore, a comparison of the value predicted by the Kriging model and the exact solution shows that its flight analysis prediction ability makes it possible to use the surrogate model in future analyses.

Assessment of the Reconstructed Aerodynamics of the Mars Science Laboratory Entry Vehicle

On 5 August 2012, the Mars Science Laboratory entry vehicle successfully entered the atmosphere of Mars, flying a guided entry until parachute deploy. The Curiosity rover landed safely in Gale crater upon completion of the entry, descent, and landing sequence. Preflight aerodynamic predictions are compared with the aerodynamic performance of the entry capsule identified from onboard flight data, including inertial-measurement-unit accelerometer and rate gyro information, and heat shield surface pressure measurements. From the onboard data, static force and moment coefficients have been extracted. These data are compared with the preflight aerodynamic database. The Mars Science Laboratory flight data represent the most complete and self-consistent record of a blunt capsule entering Mars collected to date. These data enable the separation of aerodynamic performance from atmospheric conditions. The comparisons show the Mars Science Laboratory aerodynamic characteristics have been successfully identified and resolved to an accuracy better than the aerodynamic database uncertainties used in preflight simulations. A number of small anomalies have been identified and are discussed. These data will help improve aerodynamic databases for future missions and will guide computational fluid dynamics development to improve predictions.

Comparisons of Two Wind Tunnel Pressure Databases and Partial Validation against Full-Scale Measurements

AbstractDatabase-assisted design (DAD) is an integrated methodology that calculates wind loadings and wind-induced internal forces. It can also calculate demand-to-capacity indexes for each structural member, and by checking whether they differ significantly from unity, determine the adequacy of the members’ structural design. Its practical usefulness depends on the availability of comprehensive aerodynamic databases. A public domain aerodynamic database produced in 2003 by the University of Western Ontario (UWO) is not sufficiently extensive to satisfy design needs generally encountered in practice. For this reason, the Tokyo Polytechnic University (TPU) recently developed comprehensive sets of aerodynamic databases that are publicly available and would fill large voids present in the UWO database. This paper presents comparisons of aerodynamic pressures and forces based on TPU and UWO data for low-rise buildings to help assess the extent to which the respective aerodynamic pressure measurements are comp...

Investigations on Missile Configuration Aerodynamic Characteristics for Design Optimization

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.

Comparison of Wind Pressure Coefficients with Wind Load Provisions

This study has compared the equivalent external pressure coefficients, (GCpf)eq, with 6 wind load provisions and wind tunnel test data. The wind load provisions are the ASCE 7-10, NBCC 2010, AS/NZS 2011, EN 2005, AIJ 2004 and KBC 2009. Experiment data on low-rise building have been obtained at the University of Western Ontario (UWO) to contribute to the NIST aerodynamic database [. For the experiment, a model with 1:12 of roof slope and 4.9 and 12.2m of eave height was used under open terrain conditions (See also Ref. [). (GCpf)eq was re-normalized based on the external pressure coefficients, GCpf, of ASCE 7-10. When compared to (GCpf)eq of the experiment data with 4.9m of eave height, consequently, ASCE 7-10 (81%), NBCC 2010 (84%), AS/NZS 2011 (70%), AIJ 2004 (68%) and KBC 2009 (53%) were all underestimated. Among them, KBC 2009 reveals the lowest value. On the contrary, EN2005 was overestimated with 122%. When the eave height was 12.2m, in addition, the same pattern was observed in most codes. EN2005 was slightly overestimated with 115%.

Airfoil-Performance-Degradation Prediction Based on Nondimensional Icing Parameters

A physics-based empirical correlation between icing conditions and the corresponding drag coefficient was developed for NACA 0012 airfoils, and compared to other three existing prediction methods. The correlation was developed based on experimental aerodynamic databases of iced airfoils, and derived using statistical methods. The correlation model also provides drag coefficients for varying angles of attack for a given icing condition. The calculated drag coefficients resulted in 33.40% mean absolute deviation with respect to reference data from three different experimental databases. To validate the proposed degradation model and to further extend the database for helicopter-rotor performance degradation, rotating ice-accretion experiments were conducted. Four ice shapes obtained at the NASA Icing Research Tunnel were reproduced on a 53.34-cm-chord, 1.37-m-radius NACA 0012 rotor blade at the Adverse Environment Rotor Test Stand facility. Ice-shape molding and casting techniques were introduced to capture delicate ice features, such as ice feathers. The iced-airfoil castings were tested in a dry-air wind tunnel. The drag-coefficient comparison between the proposed analytical determination method and the experimental results from both rotor ice testing and icing-wind-tunnel testing showed to be satisfactory, ranging from 5 to 25% depending on the icing condition. The effect of ice feathers on drag degradation was investigated. Ice-feather formation can account for up to 25% of the drag introduced by ice accretion before stall.

Aerodynamic database development of the ESA intermediate experimental vehicle

This work deals with the aerodynamic database development of the Intermediate Experiment Vehicle.The aerodynamic analysis, carried out for the whole flight scenario, relies on computational fluid dynamics, wind tunnel test, and engineering-based design data generated during the project phases, from rarefied flow conditions, to hypersonic continuum flow up to reach subsonic speeds regime. Therefore, the vehicle aerodynamic database covers the range of Mach number, angle of attack, sideslip and control surface deflections foreseen for the vehicle nominal re-entry. In particular, the databasing activities are developed in the light of build-up approach. This means that all aerodynamic force and moment coefficients are provided by means of a linear summation over certain number of incremental contributions such as, for example, effect of sideslip angle, aerodynamic control surface effectiveness, etc.Each force and moment coefficient is treated separately and appropriate equation is provided, in which all the pertinent contributions for obtaining the total coefficient for any selected flight conditions appear.To this aim, all the available numerical and experimental aerodynamic data are gathered in order to explicit the functional dependencies from each aerodynamic model addend through polynomial expressions obtained with the least squares method. These polynomials are function of the primary variable that drives the phenomenon whereas secondary dependencies are introduced directly into its unknown coefficients which are determined by means of best-fitting algorithms.

CFD Analysis of a Hypersonic Vehicle Powered by Triple-Module Scramjets

A numerical investigation has been carried out to study the longitudinal performance of a hypersonic airbreathing vehicle with highly integrated triple-module scramjets. CFD-Fastran is used to evaluate the aerodynamic performance of the vehicle at inlet-open scramjet unpowered mode, and a chemical reacting code ChemTur3D has been built to evaluate the propulsion performance of the triple-module engines at scramjet powered mode. The flow conditions for the calculations include variations of angle of attack at Mach 5.85 test point. The wall pressure and surface friction are integrated to calculate drag, lift and pitching moment coefficients to predict the combined aeropropulsive force and moment characteristics during engine operation. Finally, numerical results is compared with available ground test data to assess solution accuracy, and a preflight aerodynamic database of the vehicle could be built for the hypersonic flight experiments.

Comparison of Statistical Estimation Techniques for Mars Entry, Descent, and Landing Reconstruction

Flight data from an entry, descent, and landing sequence can be used to reconstruct the vehicle's trajectory, aerodynamic coefficients, and the atmospheric profile experienced by the vehicle. Past Mars missions have not contained instrumentation that would allow for the separation of uncertainties in the atmosphere and the aerodynamic database. The 2012 Mars Science Laboratory took measurements of the pressure distribution on the aeroshell forebody during entry and allows freestream atmospheric conditions to be partially observable. Methods to estimate the flight performance statistically using onboard measurements are demonstrated here through the use of simulated Mars data. A range of statistical estimators, specifically the extended Kalman filter and unscented Kalman filter, are used to demonstrate which estimator best quantifies the states and the uncertainties in the flight parameters. The techniques demonstrated herein are planned for application to the Mars Science Laboratory flight dataset.

Application of System Identification Techniques to Tejas Flight Test Data

Tejas, the Indian Light Combat aircraft, is a fly-by-wire aircraft with relaxed longitudinal stability that provides for enhanced agility, high maneuverability and performance. The aircraft is presently undergoing extensive flight test trails and the flight data gathered is being used by various design groups to evaluate aircraft systems, performance and aerodynamic characteristics. This paper gives an overview of the application of system identification techniques to Tejas flight test data for validation and update of the aircraft aerodynamic database. The aerodynamic characterization is carried out using two different approaches i) point model identification, and ii) coefficient level matching. Typical results are presented from both the approaches along with the time history plots from the flight updated aero database. The main purpose of this paper is show to how system identification techniques can lead to accurate determination of aerodynamic characteristics from flight test data.