No full-text available
To read the full-text of this research,
you can request a copy directly from the authors.
Abstract
In order to investigate the unsteady aerodynamic performance of flexible variable-sweep morphing aircraft during deformation process, an unsteady numerical simulation is proposed. The steady aerodynamic characteristics of flexible variable-sweep morphing aircraft in specific sweep angle are analyzed firstly. After that, three different variable sweep periods are selected to fulfill the unsteady numerical simulation. The impact on the aerodynamic characteristics of variable-sweep morphing aircraft is analyzed for different variable-sweep speeds. The difference of steady and unsteady aerodynamic characteristics is also analyzed, and the reason of this difference is investigated. The result demonstrates that: the real-time aerodynamic performance of flexible variable-sweep morphing aircraft can be optimized through sweep changes; the aerodynamic force difference in different variable sweep speeds is not very obvious; the difference between steady and unsteady aerodynamic force mainly caused by the wing is no more than 7% at the angle of attack 2°, and the horizontal tail's contribution can be neglected. The lift and drag coefficients of unsteady numerical simulation are larger than that of the steady numerical simulation, and the difference of pitching coefficients can be neglected also. In general, the lift and drag coefficients are larger at the low sweep angle when the angle of attack is smaller than 14°. The lift and drag coefficients are larger at the high sweep angle, when the angle of attack is bigger than that degree. The lift-drag ratio will reach to the maximum at the angle of attack 4°, and the lift-drag radio is larger at the small sweep angle. When the lift coefficient is smaller than 1.28 and the sweep is very small, the drag coefficient of the model is smaller, otherwise the drag is larger at the high sweep. Mechanism for the unsteady aerodynamic force is caused by wing additional-velocity and has nothing to do with flow field hysteresis. ©, 2015, Editorial Office of Acta Aerodynamica Sinica. All right reserved.
... [4][5] The new generation company's sliding skin wing solution USES a special flexible skin to deform the wing, changing its area by 70% [6], The SmartWing program USES a continuous, smooth, non-hinged, high-deflection high-speed control surface to improve roll and pitch performance [7], Many researchers have studied the aerodynamic characteristics and dynamic modeling of the variant aircraft. Yao et al. [8] studied the unsteady aerodynamic characteristics of the variable swept-back aircraft, and proved that the aerodynamic force caused by different swept-back speeds is not very different. Wang et al. [2] conducted the simulation of unsteady aerodynamic characteristics of a variable forward sweeping-wing aircraft, and summarized the variation rules of aerodynamic characteristics in the variation process. ...
The variable-sweep morphing aircraft (VMA) has both high speed and low speed flight performance, and can be competent for combat tasks and practical requirements in different flight environments. Aerodynamic characteristics of VMA have a great impact on flight performance. The transonic stage is particularly important in the whole variation process. It is also one of the most difficult problems in aerodynamic layout design of aircraft. In order to explore the aerodynamic characteristics of the VMA at transonic speeds, this paper adopts an efficient engineering algorithm. The aerodynamic characteristics of VMA at different Sweepback Angle, Angle of attack, Mach number and Altitude are analysed. The result demonstrates that: Sweepback Angle has great influence on aerodynamic characteristics of aircraft. For the lift coefficient, at subsonic speeds, small sweep has great influence. When the Angle of attack (AOA) is 12°, the sweepback changes from 45° to 25°, and the lift coefficient increases by 48.9%. At transonic speeds, large sweepback has great influence. When the AOA is 12°, the sweepback changes from 75° to 55°, and the lift coefficient increases by 85.7%. For the drag coefficient, at subsonic speeds, small sweepback has great influence. When the AOA is 12°, the sweepback changes from 45° to 25°, and the drag coefficient increases by 45.5%.All the changes of sweepback have great influence on drag coefficient at transonic speeds. The influence of sweepback on lift-drag ratio is complex. In general, when the Mach number increases, the sweepback corresponding to the maximum lift-drag ratio increases. The altitude factor only has some influence on the drag coefficient at subsonic speeds. These aerodynamic change rules can provide a basis for the shape design and dynamic analysis of VMA.
The Morphing Aircraft Structures (MAS) program is a Defense Advanced Research Projects Agency (DARPA) led effort to develop morphing flight vehicles capable of radical shape change in flight. Two performance parameters of interest are loiter time and dash speed as these define the persistence and responsiveness of an aircraft. The geometrical characteristics that optimize loiter time and dash speed require different geometrical planforms. Therefore, radical shape change – usually involving wing area and sweep – allows vehicle optimization across many flight regimes. The second phase of the MAS program consisted of wind tunnel tests conducted at the NASA Langley Transonic Dynamics Tunnel to demonstrate two morphing concepts and their enabling technologies with large-scale semi-span models. This paper will focus upon one of those wind tunnel tests that utilized a model developed by Lockheed Martin Aeronautics Company (LM). Wind tunnel success criteria were developed by NASA to support the DARPA program objectives. The primary focus of this paper will be the demonstration of the DARPA objectives by systematic evaluation of the wind tunnel model performance relative to the defined success criteria. This paper will also provide a description of the LM model and instrumentation, and document pertinent lessons learned. Finally, as part of the success criteria, aeroelastic characteristics of the LM derived MAS vehicle are also addressed. Evaluation of aeroelastic characteristics is the most detailed criterion investigated in this paper. While no aeroelastic instabilities were encountered as a direct result of the morphing design or components, several interesting and unexpected aeroelastic phenomenon arose during testing.
This work describes the design and construction of a fully adaptive aircraft configuration used as an experimental testbed for aerodynamic modeling and flight control. The adaptive model is designed to achieve large scale shape changes in order to investigate morphing for multi-mission UAVs. There are five independent planform changes along with independent twist control for each wing. Wind tunnel testing was conducted in Virginia Tech's Stability tunnel to analyze the aerodynamic characteristics and evaluate the usefulness of having a UAV with multiple configuration capability. Wind tunnel tests of various planform configurations indicate that different configurations yields minimum drag over a range of flight conditions.
Zero-net mass-flux jet based control of flow separation over a stalled airfoil is examined using numerical simulations. Two-dimensional simulations are carried out for a NACA 4418 airfoil at a chord Reynolds number of 40,000 and angle of attack of 18 deg. Results for the uncontrolled flow indicate the presence of three distinct natural time scales in the flow corresponding to the shear layer, separation bubble, and wake regions. The natural frequencies are used to select appropriate forcing frequencies, and it is found that forcing frequencies closer to the separation bubble frequency elicit the best response in terms of reduction of separation extent and an improvement in aerodynamic performance. In contrast, higher forcing frequencies closer to the natural shear layer frequency tend to enhance separation. The vortex dynamics and frequency response of flow are examined in detail to gain insight into mechanisms underlying the observed behavior.
Combining the characteristics of suction-transition control and micro-blowing turbulence-drag-reduction technology, a novel suction-blowing-control technique is put forward to reduce airfoil drag. By using the computational fluid dynamics (CFD) code of Fluent, the Wilcox transition mode is modified by a user-defined function (UDF) and incorporated into the CFD code. Based on the above modifications, the effect of suction-blowing control on the airfoil drag performance is numerically studied. The results show that when the suction flow rate are in a certain range, suction and suction-blowing control can both reduce the total drag of an airfoil. At the same Reynolds number, the leading-edge suction without blowing reduces the total drag below the airfoil by up to 3%, while the suction-blowing control results in a drag reduction by up to 16%. Therefore, the suction-blowing control technique is an effective way for drag reduction control.
In order to explore the unsteady aerodynamic characteristics and its mechanism of aircraft during shape morphing, wind tunnel experiments of a sliding-skin variable-sweep morphing unmanned aerial vehicle(UAV) were conducted, and unsteady aerodynamic characteristics were studied. The results demonstrate that: the unsteady aerodynamic characteristics curves of the UAV during shape morphing generate hysteresis loops around corresponding quasi-steady curves; the faster the UAV morphs, the greater the hysteresis effect is; when the UAV varies sweep with a speed of 7.5°/s, the value of the unsteady aerodynamic characteristics will deviate corresponding quasi-steady value by exceeding 5%; the main reasons of the formation of the above unsteady aerodynamic characteristics may be flow-field structure hysteresis and wing additional-velocity.
The objective of this study was to analyze the effectiveness of aerodynamic plasma actuators as a means of active flow control over an airfoil at multiple angles of attack under low Reynolds number conditions. Each angle of attack corresponded to two different flow separation mechanisms (i.e., laminar separation bubble (LSB) and fully turbulent flow separation at stall conditions). Detailed parametric studies based on steady and unsteady Reynolds Averaged Navier-Stokes simulations were performed for a NACA 0012 airfoil at a chord Reynolds number of 105 to investigate the influence of the number, the location, the imposed body force magnitude, and steady vs. unsteady operation of plasma actuators on flow control effectiveness. For LSB control, as much as a 50% improvement in the lift to drag ratio was observed. Results also show that the same improvement was achieved using unsteady or multiple actuators, which can require as much as 75% less time averaged body force compared to a single, steady actuator. For the stalled airfoil case, significant recovery in aerodynamic performance was observed for a single, steady actuator. However, for the stall conditions considered in this study, unsteady and multiple actuator configurations do not provide the same enhancement as a single, steady actuator, which may be due to the nature of the flow separation (turbulent, trailing edge separation). The results of both cases show that the optimum location of a plasma actuator would be just upstream of the separation location for maximum effectiveness. This highlights the usefulness of multiple actuator systems for flow control over a range of operating conditions as the separation location may be dynamic.
A project on smart wing program for the development and demonstration of smart materials was discussed. The concept of the program was to improve the aerodynamic and aeroelastic performance of military aircrafts. The program was divided into two phases which included the development of adaptive wing structures with integrated actuation mechanism to replace standard hinged control surface and in the second phase, wind tunnel test was performed which provided baseline steady state data at Mach numbers ranging from 0.3 to 0.8. The work performed under the program demonstrated the feasibility of developing smart control surface designs for optimal aerodynamic performance.
An experimental study of flow control using an array of three synthetic jets has been undertaken in a separated laminar flow over an inclined flat plate in a water channel. Particle image velocimetry was employed to obtain the information about the extent of flow separation delay at different synthetic jet operating conditions. A laser-induced florescence flow visualization technique was used to reveal the characteristics of the vortical structures produced by the synthetic jets, which result in a delay of separation. It was observed that the flow separation delay is typically associated with the presence of two or three streaks of high streamwise velocity in the separated flow. Based on the results from the present experiment, a parameter map indicating the flow patterns observed at different synthetic jet operating conditions is produced. In addition, a contour map showing the separation control effectiveness at the corresponding conditions is also obtained. It was found that the two-streak flowpattern is mostly produced at a jet-to freestream velocity ratio between 0.3 and 0.5, whereas the three-streak pattern occurs at a velocity ratio between 0.5 and 1.5. The laser-induced florescence images confirm that the two-streak flow pattern is caused by the hairpin type vortical structures produced by synthetic jets, whereas the three-streak pattern is caused by the tilted vortex ring type structures. For the range of actuator operating conditions tested in this study, operating the synthetic jets at a dimensionless stroke length of around 2, a velocity ratio of around 0.5, and a Strouhal number of around 1.6 would deliver the best flow control effect with the least energy consumption. Under these conditions, hairpinlike vortical structures are observed with a streamwise spacing of 44% of the local boundary-layer thickness at the orifices of the synthetic jets. In this experiment, the spacing between the jets and the distance between the jet array and the baseline separation line are fixed. As the level of interaction between neighboring synthetic jets and their flow control effectiveness are expected to alter upon changes in these two parameters, the generality of this finding remains to be established in future work.
Boundary-conforming coordinate transformations are used widely to map a
flow region onto a computational space in which a finite-difference
solution to the differential flow conservation laws is carried out. This
method entails difficulties with boundary conditions, with maintenance
of global conservation, and with computation of the local volume element
under time-dependent mappings that result from boundary motion. To
improve the method, a differential 'Geometric Conservation Law' (GCL) is
formulated that governs the spatial volume element under an arbitrary
mapping. The GCL is solved numerically along with the flow conservation
laws using conservative difference operators. Numerical results are
presented for both implicit and explicit difference schemes applied to
the unsteady Navier-Stokes equations and to the steady supersonic flow
equations.
Morphing Aircraft Concepts, Classifications, and Challenges Industrial and Commercial Applications of Smart Structures Technologies
- A K Jha
- J Kudva
Jha A K, Kudva J N. Morphing Aircraft Concepts, Classifications,
and Challenges[C]//Anderson E H. Smart Structures and Materials
2004: Industrial and Commercial Applications of Smart Structures
Technologies. Bellingham:SPIE, 2004: 213-224.
[2]
Mission Effectiveness Comparisons of Morphing and Non-Morphing Vehicles A Study of the Benefits of Using Morphing Wing Technology in Fighter Aircraft Systems Control of Shock-Boundary Layer Interactions (SBLIs) Using An Oscillatory Jet[R]
- J C Bowman
- R W Plumley
- J A Dubois
Bowman J C, Plumley R W, Dubois J A, et al. Mission Effectiveness
Comparisons of Morphing and Non-Morphing Vehicles[R]. AIAA
2006-7771, 2006.
[3]
Smith K, Butt J, Von Spakovsky M R. A Study of the Benefits of
Using Morphing Wing Technology in Fighter Aircraft Systems[R].
AIAA 2007-4616, 2007.
[4]
Hassan A A, Osborne B, Schwimley S, et al. Control of
Shock-Boundary Layer Interactions (SBLIs) Using An Oscillatory
Jet[R]. AIAA 2007-476, 2007.
[5]
Raju R, Mittal R, Cattafesta L. Dynamics of Airfoil Separation
Control Using Zero-Net Mass-Flux Forcing[J]. AIAA Journal, 2008,
46(12): 3103-3115.
[6]
An experimental investigation of flow separation control sing an array of synthetic jets
- S Y Zhang
- S Zhong
- 段会申
- 何雨薇 刘沛清
- 等 二维翼型微吸吹气减阻控制新技术
- 数值研究
Zhang S Y, Zhong S. An experimental investigation of flow
separation control sing an array of synthetic jets[R]. AIAA 2009-4185,
2009.
[7]
段会申, 刘沛清, 何雨薇, 等. 二维翼型微吸吹气减阻控制新技术
数值研究[J]. 航空学报, 2009, 30(7): 1219-1226.
Executive Summary AFTI/F-111 Mission Adaptive Wing
- J Hall
Hall J M. Executive Summary AFTI/F-111 Mission Adaptive
Wing[R].
Compendium of unsteady aerodynamic measurements
- Ara R H Landon
R H Landon, ARA. Compendium of unsteady aerodynamic
measurements[R].