Fighter Wing Research Articles

Computational Analysis of Fighter Aircraft Wing under Mach Number 0.7 for Small Sweep Angles

Fighter aircraft wings are the leading lift generating components for any aerospace vehicle. The recital of any flying vehicle largely depends on its wing design. Missiles and the fighter aircrafts which are having propulsion system mostly have fins to control and maneuver. In this present paper work an attempt has been made to design a fighter aircraft wing configuration which will be used in some air launched air to surface guided weapons fighter aircraft. The main focus of this paper agreement in determining the Sweep-back effects on fighter aircraft wing under transonic condition at different angles of attack (AoA) from 0 to 5 degrees. For this the fighter aircraft wing performance for various flow conditions and sweep angles are obtained based on the empirical, semi-empirical and CFD simulation results. Hence by studying these computational results would help in the optimizing geometry for better performance, an finest wing design for the air launched air to surface body with conservative wing can be obtained.

Wing Structure Static Analysis using Superelement

This paper describes the use of superelement for the stress and deflection analysis of a typical fighter wing structure. Three methods of analyses were carried out and compared: practical/theoretical analysis; finite element analysis with the conventional element modeling approach; and finite element analysis with the superelement modeling. Finite element models of the wing were developed. NASTRAN FE software was used for finite element analysis. CQUAD4 and BAR2 elements represent the individual components of the wing such as the skin and stringers. In superelement approach, the wing was divided to four (4) substructures known as superelements. Wing loading at 1-g flight condition was assumed for analysis. For all these methods, the direct stress and deflection are sought and to be compared. Result shows a good agreement between the three methods.

Wing-Box Weight Optimization Using Curvilinear Spars and Ribs (SpaRibs)

The aim of this research is to perform topology and sizing optimization of wing-box structures using curvilinear spars and ribs, referred to as SpaRibs in the following. To accomplish this, a new framework calledEBF3SSWingOpt is being developed at Virginia Polytechnic Institute and State University. The optimization framework includes two different methodologies: a one-step optimization methodology where topology and sizing optimization are carried out together and a two-step optimization methodology where topology and sizing optimization are carried out separately using different constraints and objective functions. A description of how the general framework is developed and applied for optimizing winglike structures is provided and the optimization-problem formulation is stated. Two practical design problems solved using EBF3SSWingOpt are presented: a rectangular wing box and a generic fighter wing. In both cases, the structure with the SpaRibs is lighter than the initial structure with straight spars and ribs.Moreover, the two-step optimization framework has proven to be better atfinding anoptimal solution than the one-step framework. Finally, different designs with comparable weights but different stress distributions, buckling properties, and dynamic behaviors were found.

The quantification of fatigue crack initiators in aluminium alloy 7050‐T7451 using quantitative fractography

PurposeThis paper aims to present some aspects associated with the life prediction of structures with fatigue cracks growing from small natural discontinuities in aluminium alloy (AA)7050‐T7451 for a surface condition that is present in F/A‐18 A/B aircraft critical structure.Design/methodology/approachFatigue results are presented for thick section AA7050 plate coupons loaded with a representative fighter aircraft wing root bending moment loading spectrum. Detailed quantitative fractography (QF) was used to gain a deeper understanding of issues relevant to an improved fatigue life predictive capacity for this material by using the QF results to investigate the “effectiveness” of the fatigue initiating discontinuities.FindingsEstimates of the “effectiveness” of the fatigue initiating discontinuities as quasi pre‐existing fatigue cracks (“equivalent pre‐crack size” (EPS) here) were made with the aid of a simple crack growth model. This model, based on experience, was found to be valid for the applied spectrum and stress levels used. These stress levels were chosen to represent those that may be found in highly stressed locations of fighter aircraft; and as such would usually lead to the limiting fatigue life of such a structure.Research limitations/implicationsThe method has been extended to other crack growth situations and is being used to build a database large enough to determine the best probability distribution of the “effectiveness” of the fatigue initiating discontinuities for not only the surface condition reported here but several other surface conditions typical of aircraft metallic structure.Practical implicationsThe EPS of the discontinuities from which the cracks grew were used to investigate distributions that may be used in a risk‐based assessment using deterministic crack growth measurements from such discontinuities. Some of the problems that remain to be resolved in such an analysis, prior to its use in a risk‐based assessment are discussed.Originality/valueThis work improves the understanding of the interaction of small fatigue cracks generated by representative loading spectra with the small discontinuities from which they grow and shows that the fatigue process is remarkably consistent down to very small sizes.

Reduced order model of three-dimensional Euler equations using proper orthogonal decomposition basis

This study seeks to validate the accuracy and the efficiency of the aerodynamic reduced order model (ROM). In doing this, snapshot data are generated from the full system analysis of a fighter wing problem. From an eigensystem analysis of these snapshots, the basis vector reproducing the behavior of the full system is obtained. The span length, sweep angle, dihedral angle, and spar and rib thickness representing the wing configuration are determined as the input variables. The constructed ROM is applied to the fighter wing problem while varying the input conditions for validation. Subsequently, a comparison of the reduced system with the full system confirmed that the aerodynamic performance is within 4% error and that the L2 norms are 10−6 order of the entire flow field. Therefore, the ROM is able to capture the variation of the aerodynamic performance with respect to the input variables. Though there are structural input variables which influence the aerodynamic performance indirectly, the ROM can reproduce the flow field of the full system. Additionally, even if the ROM incurs a high computational cost to generate snapshots, it can represent the behavior of the full system efficiently once the reduced order model is constructed.

Analytical study of fatigue crack growth in AA7050 notched specimens under spectrum loading

Abstract An analytical study of fatigue crack growth in aluminium alloy 7050-T7451 notched specimens under a fighter aircraft wing root bending moment spectrum was conducted. The crack growth data were measured by quantitative fractography for three groups of specimens with different stress concentration geometrical features. Under spectrum loading and for each spectrum peak stress level, a minimum of five specimens were tested. Based on the analysis of the measured spectrum crack growth data using linear elastic fracture mechanics, it was found that the concept of geometry factors formulated in the stress intensity factor could not collapse the crack growth rate data derived from each stress concentration feature, particularly near the small crack growth region. In order to investigate the possible reasons for this, three-dimensional elastic–plastic finite element analysis was used to determine notch plastic zone sizes for each stress concentration geometry. As a consequence, an alternative crack growth driving force by considering both notch elastic–plastic stress field and gross net-section stress field was proposed and used to interpret the fatigue crack growth data under spectrum loading. It was found that the predictions of crack growth under spectrum loading for different stress concentration factors at different peak load levels agree reasonably well with the experimental results.

신경망으로 구축된 불확실성 모델을 이용한 전투기 날개의 강건 최적 설계

This study performed robust design optimization of fighter wing planform, considering uncertainty based on neural network model. To construct uncertainty model, aerodynamic performance and their sensitivity were evaluated by 3-dimensional Euler equations and adjoint variable method at experimental points selected from central composite design. In addition, because a neural network model has the advantage of capturing non-linear characteristic, it was possible to predict sensitivity of the aerodynamic performance efficiently and accurately . From the results of robust design optimization, it could be confirmed that the robustness of the objective function and constraints were improved if the variation of uncertainty and sigma level were increased.ࠀȀ耀Ѐ€

Predicting the Cost Per Flying Hour for the F-16 Using Programmatic and Operational Data

Abstract We analyzed economic, operational, and programmatic data from Air National Guard and active duty F-16C/D fighter wings in search of explanatory variables that influence a wing's cost per flying hour (CPFH). Using data from the Air Force Total Ownership Cost database and from the Air Force Knowledge Systems database, we evaluated the predictive ability of various operational and programmatic variables, such as aircraft age, average sortie duration, base location, utilization rate, percent engine type, percent block, and previous year's CPFH. Although separate models were required for active duty and Air National Guard units, both regression models incorporated utilization rate, base location, percent block, and percent engine type and appeared to accurately predict an F-16 C/D fighter wing's CPFH.

충격응답함수와 조정법을 이용한 항공기 날개의 충격하중 복원 연구

충격응답함수와 조정법(regularization methods)을 이용하여 항공기 날개의 충격하중 복원 가능성을 검토하였다. 충격하중에 대한 구조의 응답을 내타낼 수 있는 충격응답함수를 날개 유한요소모델의 강성과 질량 자료로 유도하였다. 일반적으로 부적합(ill-posed) 특성을 지닌 충격응답함수의 역행렬은 반복 Tikhonov 조종법(Iterative Tikhonov Regularization Method)과 일반화 특이치 분해법(Generalized Singular Value Decomposition Method)을 사용하여 구하였다. 수치적 입증을 위하여 전투기급 주익을 사용하였다. 해당 주익의 유한요소해석을 통하여 임의의 충격하중에 대한 변위와 변형률을 계산하였으며, 이를 충격응답함수로 계산한 결과와 비교하였다. 또한, 유한요소해석에서 계산된 변형률을 사용하여 충격하중을 복원하였다. 수치적 입증 결과 항공기 구조의 충격하중 모니터링이 본 방법으로 가능할 수 있음을 보여주었다. The capability for reconstructing impact forces of an aircraft wing using impact response functions and regularization methods were examined. The impact response function which expresses the relation between the structure response and the impact force was derived using the information on mass and stiffness data of a finite element model for the wing. Iterative Tikhonov regularization method and generalized singular value decomposition method were used to inverse the impact response function that was generally ill-posed. For the numerical verification, a fighter aircraft wing was used. Strain and deflection histories obtained from finite element analysis were compared with the results calculated using impact response functions. And the impact forces were reconstructed with the strain histories obtained from finite element analysis. The numerical verification results showed that this method can be used to monitor impact forces on aircraft structures.

Multi-Objective and Multidisciplinary Design Optimization of Supersonic Fighter Wing

An aerodynamic/structural multidisciplinary design with multiple objectives was carried out for the supersonic fighter wing using response surface methodology. Through a series of static aeroelastic analyses of a variety of candidate wings, the aerodynamic performance and structural strength were calculated. Nine wing and airfoil parameters were chosen for the aerodynamic design variables, and four structural variables were added to determine the wing skin thickness. To consider various flight conditions, multipoint design optimization was performed on the three representative design points. As expected, the single-point design shows the most improved performance on its own design point, but it produces inferior results by not satisfying some constraints on other design points. To improve the performances evenly and moderately at all design points, a multipoint optimal design was conducted. A genetic algorithm was also introduced to control the weight of the multiple objectives. The multipoint designed wing features improved performance and satisfies whole constraints at all design points. It is similar to the real supersonic fighter wing that was developed through numerous wind-tunnel tests and tradeoff studies. The proposed multidisciplinary design optimization framework could be adopted as an efficient practical design tool for the supersonic fighter wing to fix the basic geometry at a conceptual design stage.

An Integrated Management Approach for Nesting Osprey to Protect Human Safety and Aircraft at Langley AFB, Virginia

Author(s): Olexa, Thomas Joseph | Abstract: North American osprey are increasingly becoming a serious aviation safety concern to both military and civilian aircraft. Since 1985, the United States Air Force documented 25 osprey strikes with aircraft, resulting in excess of $1 million dollars in damage. In 50% of the osprey strikes reported to the National Wildlife Strike Database, the aircraft was damaged. Osprey are present from March through September at Langley Air Force Base, where more than 2 dozen nesting pairs have been identified on or immediately adjacent to the airfield. The habituated nesting and breeding behavior of osprey at Langley predisposes this species to impacting aircraft arriving or departing the airfield. The apparent strike risk and flight safety concerns associated with nesting osprey resulted in the development of an integrated hazard/damage management program. As part of the 1st Fighter Wing Bird Aircraft Strike Hazard Plan, the program incorporated a diverse management approach that included nest surveys, behavior monitoring, exclusionary practices, nest removals, egg oiling, traditional hazing, lethal reinforcement, and nestling translocation. Preliminary analysis of the project (2000-2004) suggests the number of nest sites remain constant (range 26-36) from year to year; however, airfield use by osprey has declined 62% since the inception of the program. Exclusion practices, egg oiling, and juvenile translocation are presumed to be the most effective strategies in discouraging nesting and reducing airfield occurrences. Traditional hazing and nest removals had no effect on discouraging osprey behavior or nest site preference. This program will continue to integrate and evaluate management techniques for resolving aviation and human conflicts associated with nesting osprey and may ultimately serve as a technical guide for professional wildlife damage agencies, aviation safety personnel, and natural resource managers.

MDO 프레임워크 개발을 통한 항공기 날개 통합최적화 설계

다분야 통합 최적화 설계 환경을 제공하는 소프트웨어 체계인 MDO 프레임워크 개발을 위해서는 다양한 운영체계와 언어에서 개발된 해석코드들의 통합, CAD 및 데이터베이스 시스템과의 통합, 복잡한 GUI 환경의 구현 등이 필수적으로 요구되고 해석코드의 추가, 새로운 MDO 기법의 도입에 따른 수정 및 확장에 대한 고려가 충분히 이루어져야 한다. 본 논문에서는 MDO 프레임워크의 설계단계부터 고려되어야 할 사항들과 각 구성요소들의 시스템 통합 방법을 연구, 적용 방안을 제시하며 이를 바탕으로 비행체 통합 최적설계 시스템 환경을 구현하였다. MDF 및 CO 기법 등 대표적인 MDO 기법을 적용할 수 있는 데이터베이스 설계과정을 정립하고, 구현된 통합 최적설계 시스템을 이용하여 전투기 날개 형상 최적 설계를 수행하여 개발된 MDO 프레임워크의 효율성 및 유용성을 검증하였다. 구배 기반 최적화 기법을 이용하여 삼십번의 설계 반복으로 최적 날개 형상을 도출하였다. MDO framework, which provides multidisciplinary system design and optimization environment, requires integration of the analyses codes developed at various computer languages and operating systems, integration of CAD and DBMS, and development of complex GUI. Emphases must be given to the software modification and upgrades in conjunction with the analysis code addition and MDO method implementation. In this study, techniques about system integration and analysis code interface have been studied extensively, and the database design and communication methods which can handle the MDO methods like MDF and CO have been studied. Using the dedicated MDO framework developed for the air vehicle design, the multidisciplinary fighter aircraft wing design has been performed to demonstrate the efficiency and usefulness of the software. Optimum wing configuration is derived using the gradient-based optimization methods within thirty design iterations.

Aerodynamic Performance of the Smart Wing Control Effectors

This paper presents and discusses the wind tunnel results obtained during Phase 1 and Phase 2 of the Smart Wing program. A series of four tests was conducted, two in each phase. The purpose of the tests was to evaluate the performance of smart material based control effectors in representative aerodynamic environments. In Phase 1, wing twist using a Shape Memory Alloy (SMA) torque tube and smoothly contoured trailing edge surfaces enabled by SMA wires were evaluated on a semi-span, 16% scale model of a typical fighter wing. The focus of Phase 2 was to address the design and demonstration of a high-frequency, large-deflection, and smoothly contoured trailing edge control surface capable of a spanwise variation in deflection. The actuation system for this control surface was based on piezoelectric (PZT) ultrasonic motors. A smoothly contoured leading edge control surface was also built and tested using SMA wires. These designs were evaluated on a 30%, full-span model of a representative Unmanned Combat Air Vehicle (UCAV). In each wind tunnel entry comparisons were made to the performance of conventional control surfaces. Successful results included: improved aileron effectiveness at high dynamic pressures; demonstrated improvements in lateral and longitudinal control effectiveness with smoothly contoured control surfaces over conventional hinged control surfaces; continuous spanwise shape control; and, large deflections at rates over 80°/s, which is well within the desired deflection aileron rates for control of fighter aircraft.

Control Surface Reversal in the Transonic Regime Including Viscous Effects

This paper investigates control surface reversal in the transonic flight regime. Surface-pressure distributions on a rigid and deformable wing are presented using linear and nonlinear aerodynamic methods, which incorporate the effects of surface shocks and viscosity with an interactive boundary layer. Transonic small-disturbance (TSD) aerodynamics is used to determine reversal of a typical fighter wing to study the interactions among control surface deflections, structural flexibility, and embedded shocks in a flowfield where significant viscous effects can exist. Pressure distributions on the wing are examined, and control surface reversal calculations are presented. The results show that the reduction of the reversal dynamic pressure can be as much as 30% when an aerodynamic method capable of representing a shock is used. A limited number of Euler computational fluid dynamic (CFD) results are presented for the purpose of comparison with transonic small disturbance results. The adequacy of using TSD theory in reversal prediction as compared to that obtained from either Euler or Navier-Stokes CFD aerodynamics is discussed. It is suggested that the inviscid transonic small disturbance aerodynamic provides a capability of representing shocks with only a modest increase of computer time beyond that necessary with the linear surface panel method. It is recommended that the inviscid version of TSD be incorporated in the preliminary design of aircraft operating at transonic speeds.

Assessing unit readiness of an air force fighter wing

(2000). Assessing unit readiness of an air force fighter wing. Strategic Analysis: Vol. 24, No. 4, pp. 695-706.

Computation of unsteady transonic flow over a fighter wing using a zonal Navier–Stokes/full‐potential method

An improved hybrid method for computing unsteady compressible viscous flows is presented. This method divides the computational domain into two zones. In the inner zone, the Navier–Stokes equations are solved using a diagonal form of an alternating-direction implicit (ADI) approximate factorisation procedure. In the outer zone, the unsteady full-potential equation (FPE) is solved. The two zones are tightly coupled so that steady and unsteady flows may be efficiently solved. Characteristic-based viscous/inviscid interface boundary conditions are employed to avoid spurious reflections at that interface. The resulting CPU times are about 60% of the full Navier–Stokes CPU times for unsteady flows in non-vector processing machines. Applications of the method are presented for a F-5 wing in steady and unsteady transonic flows. Steady surface pressures are in very good agreement with experimental data and are essentially identical to the full Navier–Stokes predictions. Density contours show that shocks cross the viscous/inviscid interface smoothly, so that the accuracy of full Navier–Stokes equations can be retained with significant savings in computational time. Copyright © 1999 John Wiley & Sons, Ltd.

Computation of unsteady transonic flow over a fighter wing using a zonal Navier-Stokes/full-potential method

An improved hybrid method for computing unsteady compressible viscous flows is presented. This method divides the computational domain into two zones. In the inner zone, the Navier-Stokes equations are solved using a diagonal form of an alternating-direction implicit (ADI) approximate factorisation procedure. In the outer zone, the unsteady full-potential equation (FPE) is solved. The two zones are tightly coupled so that steady and unsteady flows may be efficiently solved. Characteristic-based viscous/inviscid interface boundary conditions are employed to avoid spurious reflections at that interface. The resulting CPU times are about 60% of the full Navier-Stokes CPU times for unsteady flows in non-vector processing machines. Applications of the method are presented for a F-5 wing in steady and unsteady transonic flows

Multiple order pole pure lag rational function approximations for unsteady aerodynamics

Most existing rational function approximations for the time domain representation of unsteady generalized airloads lead to an ill-conditioned system dynamic matrix in the presence of closely spaced poles. A new class of multiple order pole pure lag rational function approximations (RFA) is presented in this article to overcome this problem. The present class of approximation s is developed as a consistent generalization of an existing simple pole pure lag RFA while preserving the resulting state vector dimension. A nonlinear nongradient optimization technique is used for the computation of the lag poles in the approximation. The structure of the proposed class of approximations preserves the pure lag form, thus allowing for specific physical interpretations of the individual terms in the approximation and renders the subsequent optimization problem simpler, in that fewer constraints need to be imposed during the optimization process. Furthermore, the new class of approximations leads to substantial reduction in computational costs for optimization for a given fit accuracy, in comparison to existing simple pole approximations. Results are presented for the case of a typical delta wing fighter aircraft configuration.

Application of Navier-Stokes aeroelastic methods to improve fighter wing maneuver performance

An aeroelastic analysis method, based on three-dimensional Navier-Stokes aerodynamics, has been applied to improve the performance of fighter wings operating at sustained maneuver flight conditions. The scheme reduces the trimmed pressure drag of wings performing high-g maneuvers through a simultaneous application of control surface deflection and aeroelastic twist. The aerodynamic and structural interactions are decoupled by assuming an aeroelastic twist mode shape and optimizing the aerodynamic performance based on this aeroelastic mode. The wing structural stiffness properties are then determined through an inverse scheme based on the aerodynamic loads and desired twist at the maneuver flight condition. The decoupled technique is verified by performing a fully coupled aeroelastic analysis. One of the more important features of this application, over and above improved maneuver flight performance, is that the wing performance at cruise conditions is not compromised. Thus, this method represents a multiple-point wing design capability utilizing computational aerodynamics methods and aeroelastic tailoring.

Ten Leadership Steps to a Smooth Facility Upgrade

Leading a medical treatment facility (MTF) through the construction phases of a major upgrade, repair, or any sizable facility project is a challenging and complex endeavor. The Executive Staff can make the process easier on everyone and ensure the final product is in tune with the MTFs needs by taking positive leadership actions. This article describes the effective actions discovered during a 2.5-year-long total MTF upgrade of the 52 Tactical Fighter Wing (TFW) Clinic. These actions involved promulgating a broad systems view of the process, establishing philosophies and criteria, promoting team work, involving customers, and building consensus among the staff.