Delta Wing Model Research Articles

The leading-edge vortex of swift wing-shaped delta wings.

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

Aeroelastic flutter analysis considering modeling uncertainties

A method for efficient flutter analysis of aeroelastic systems including modeling uncertainties is presented. The aerodynamic model is approximated by a piece-wise continuous rational polynomial function, allowing the flutter equation to be formulated as a set of piece-wise linear eigenproblems. Feasible sets for eigenvalue variations caused by combinations of modeling uncertainties are computed with an approach based on eigenvalue differentials and Minkowski sums. The method allows a general linear formulation for the nominal system model as well as for the uncertainty description and is thus straightforwardly applicable to linearized aeroelastic models including both structural and aerodynamic uncertainties. It has favorable computational properties and, for a wide range of uncertainty descriptions, feasible sets can be computed in output polynomial time. The method is applied to analyze the flutter characteristics of a delta wing model. It is found that both structural and aerodynamic uncertainties can have a considerable effect on the damping trends of the flutter modes and thus need to be accounted for in order to obtain reliable predictions of the flutter characteristics. This indicates that it can be beneficial to allow a flexible and detailed formulation for both aerodynamic and structural uncertainties, as is possible with the present system formulation.

Computational and Analytical Investigation of Aerodynamic Derivatives of Similitude Delta Wing Model at Hypersonic Speeds

This research paper presents a computational and analytical investigation of aerodynamic derivatives in an oscillating wedge. Unsteady hypersonic similitude has been apprehended for an oscillating wedge with an attached bow shock at a large incidence angle. The problems of instability and shock waves are generally associated with hypersonic flow and, therefore, it is imperative to evaluate aerodynamic models that can solve these problems. Lighthill’s piston theory is an unsteady aerodynamic model that is valid for an oscillating wedge with an attached shock wave. The analytical solution verifies that both the stiffness and the damping derivatives attain high values when the semi-vertex angle of the wedge is increased, while both derivatives assume lower values at increasing Mach numbers. Similarly, the pressure distribution over the wedge is evaluated to determine the details of how the developing flow cause the instabilities. Our study presents the contour plots of pressure, temperature, density, and Mach number that unravels the positions of flow separations in an oscillating wedge model.

Aerodynamic Flow Characteristics of Utilizing Delta Wing Configurations in Supersonic and Subsonic Flight Regimes

Computational fluid dynamic tests are performed on delta wing models at different heights and speeds in order to achieve lift and drag coefficient values. Primarily, testing was done at supersonic speeds to reveal the advantages of these wing configurations at supersonic flight regimes at a cruise speed and altitude. The low speed characteristics are also examined, important for take-off and landing regimes where the distinctive vortices become prominent. Throughout the two flight conditions tested, a simple delta wing model (with a straight swept wing) is compared to a delta wing model that exhibited a ‘LERX’ (leading edge root extension). Provided literature describes how the performance of delta wings can be improved through this inclusion. Results obtained from the tests show that the model with the LERX has a small, but significant, performance improvement over the simple delta model, in respect to the maximum achievable lift coefficient and maximum stall angle. Lift to drag ratio is not improved however, due to the large vortices creating pressure drag. Generally, the delta wing models produce relatively small amounts of drag, and slightly less lower lift, when at low angles of attack. This is primarily due to the geometry of the models that have thin leading edges and also low thickness to chord ratios.

Effect of Dorsal Fin on Symmetry of Vortices over Slender Delta Wing

Three different sets of experiments are designed to reveal the behavior of the vortex flow over a delta wing with different fin heights. A sharp-edged flat-plate delta wing with an 82.6-deg-sweep angle is tested in low-speed wind tunnels at angles of attack up to 35 deg. The same tests are performed on an identical delta-wing model but with a flat-plate dorsal fin. The ratio of the local fin height to the local wing semispan varies from 0.3 to 1.5. The results of smoke laser-sheet visualization, particle image velocimetry, and force measurement are obtained and presented. The test results show that the vortices over the delta-wing-alone model at high angles of attack and zero sideslip are stable, symmetric, and steady before vortex breakdown, but a low dorsal fin renders the originally symmetric vortices asymmetric at the same angle of attack; the flow recovers symmetry and conicity when a higher fin is added. These results are in agreement with predictions by the stability theoryCaiet al., (“Stability of Symmetric Vortices in Two Dimensions and over Three-Dimensional Slender Conical Bodies,” Journal of Fluid Mechanics, Vol. null, 2003, pp. 65–94). This agreement demonstrates that the absolute type of instability is responsible for loss of symmetry of the vortices over a delta wing with a fin within the range of tested conditions.

The Analysis about External Compressible Flow around an Aircraft Wing

The Steady flow simulation to selected the delta wing model for different angles of attack in the Maher number.The law of flow field changes with the angle of attack is gotten.Through the FLUENT simulation,The variation tendency of coefficient of lift and drag in the different angle of attack is gotten.Further reveals the change rule of Maher number, pressure, velocity and other parameters in the different angle of attack.With increasing angle of attack, Maher number distribution is sparse of the same position increases and the greater numerical.the distribution of velocity vector is sparse of the same position increases and the greater numerical.the pressure distribution is sparse of the same position increases and the greater numerical.

On the Effect of the Incident Mach’s Wave on the Pulsation Level in Boundary Layer of the Plane Delta Wing

The experimental results of high level fluctuation excitation by external Mach’s wave in the boundary layer of delta wing model with blunt leading edges at Mach numbers M = 2, 2.5, 4 are presented. The exitation areas and mass flow pulsation levels in the conditions of subsonic, sonic and supersonic leading edges have been defined. It was found that the maximum value of the pulsations is 12–15 % and varies only slightly from the flow conditions around of the delta wing

Leading-Edge Radius Effects on Aerodynamic Characteristics of 50-Degree Delta Wings

The experimental study focuses on the effects of the leading-edge radius on the flow over 50 deg swept delta wing models. Three models were tested; one model has a sharp leading edge, and two others have a semicircular leading edge of a different radius. Six-component balancemeasurementswere performedand indicated that the leading-edge radius affects the level and slope of the forces acting on the models. The vortical flow on and off the surface of the models was investigated using an oil flow visualization and a stereo particle image velocimetry technique. The leading-edge radius is shown to affect the location, size, and strength of the vortices and also the vortex breakdown location over the models. The flow data are correlated with the balance data.

Flow Visualization and CFD Simulation on 65° Delta Wing at Subsonic Condition

The primary objective of this study is to capture the skin friction lines as the impression of primary and secondary vortices on the upper surface of a slender sharp edged 65¡ delta wing model at low speed subsonic static test condition using oil flow visualization. Subsequently CFD (computational fluid dynamics) analysis is carried out using CFD code Ansys Fluent. Multi block structured meshes are generated within the computational domain and Reynolds Averaged Navier Stokes (RANS) based steady state three dimensional CFD simulations are performed. The experimental and computational results are compared both qualitatively and quantitatively in the context of surface skin friction topography and vortex breakdown.

Delta Wing-Fuselage Interactions - Experimental Study

Experiments were conducted on thin delta wings to investigate, for subsonic flow, the effect of both privileged apex angle values and the wing-fuselage interactions on the aerodynamic characteristics, i.e. the distribution of the defect pressure on the extrados, the drag and the lift coefficients. For this purpose, several delta wing models of various apex angle (β = 75, 80 and 85°) were realized and tested without and with fuselages of cylindrical form, with diameters of 20 and 30 mm, downstream the apex and appropriately disposed on the extrados. The impact of the apex angle as well as the interaction on the defect pressure were specially considered along the apex vortices where the pressure defect is usually maximum. The above mentioned effects were investigated via the variations of the mean velocity in the wind tunnel and the incidence (attack) angle.

EFFECT OF DORSAL FIN ON THE STABILITY OF VORTICES OVER A DELTA WING

Cai et al.5 developed a vortex stability theory for slender conical bodies and analyzed the stability of vortex pairs over slender conical wing-body combinations under small perturbations. An experimental study is presented in this paper to verify the validity of the theoretical predictions. A sharp-edged flat-plate delta wing is tested in a low-speed wind tunnel. A smoke-laser-sheet visualization technique is used to visualize and measure the positions of the vortex pair, which are found to be symmetric and conical over the wing. The same tests are performed on an identical delta wing model but with a flat-plate dorsal fin mounted vertically in the incidence plane of the wing. Two fin heights are tested. The ratios of the local fin height to the local wing semi-span are 0.75 and 1.50. The test results clearly indicate that the vortices become asymmetric and non-conical over the model with the fin height ratio of 0.75 and recover symmetry and conicity over the model with the fin height ratio of 1.50, providing direct experimental evidence of the theoretical predictions.

Prediction of F-16XL Flight-Flow Physics

T HIS special section is the result of fruitful endeavors by an international group of researchers in industry, government laboratories, and university-led efforts to improve the technology readiness level of their CFD solvers through comparisons with flight data collected on the F-16XL-1 aircraft at a variety of test conditions. These 1996flight datawere documented in [1] and detailed theflightflow physics of this aircraft through surface tufts, pressures, boundary-layer rakes, and skin-friction measurements. The flight project was called the Cranked-Arrow Wing Aerodynamics Project (CAWAP), due to its leading-edge sweep crank (70 deg inboard, 50 deg outboard), and served as a basis for the international comparisons to bemade, called CAWAPI. This highly focused effort was one of two vortical flow studies facilitated by the NATO Research and Technology Organization (RTO) through its Applied Vehicle Technology (AVT) Panel with a title of “Understanding and Modeling Vortical Flows to Improve the Technology Readiness Level for Military Aircraft.” It was given a task group number of AVT-113 and had an official start date of spring 2003. The companion part of this task group dealt with fundamentals of vortical flow from both an experimental and numerical perspective on an analytically describable delta-wing model, for which much surface pressure data had already beenmeasured at NASALangleyResearch Center at a variety of Mach and Reynolds numbers, and is called the “Vortex FlowExperiment–2” (VFE-2) as proposed in [2]. These two parts or facets helped one another in understanding the predictions and data that had been or were being collected. The CAWAPI facet had the distinction of using actual aircraft geometry, which is subject to the ITAR controls, and required much cooperation between NASA Langley Research Center, NASA Headquarters, and the leadership of the various NATO or Partnersfor-Peace participating organizations in establishing how the geometry and grids could be shared. This and other background information is contained in the first paper that follows. The second paper discusses how the supplied geometry was processed into acceptable computational grids for both the structured and unstructured solver communities, papers three to five detail the comparisons of three classes of grid solutionswith theflight data, and the last paper provides what has been learned from CAWAPI. The author is extremely proudof themany outstanding researchers and organizations that have had a part in the CAWAPI facet. These RTO task groups do not come with funded support, so each participating organization had to anticipate that the benefit it accrues would be greater than the expense encountered. The participating researchers accommodated the work often on their own time under the benevolent eye of their employers who saw this work to be of importance and provided an opportunity for their staff to make a novel contribution, as well as to test their own solvers. We have also benefited from significant graduate student involvement; in particular, CAWAPI has lasted long enough for some to complete their advanced degrees and to be part of the coauthorship of two of the following papers. The author alsowishes to thank theAIAA for providing ameans of rapid dissemination of results obtained duringCAWAPI through two special sessions at the 45th AIAA Aerospace Sciences Meeting and Exhibit (January 2007), and now through the Journal of Aircraft. In particular, Professor Frank Coton (University of Glasgow) and the Applied Aerodynamics Technical Committee are thanked for advocating and facilitating these special sessions at the general meeting in which results from nine organizations were reported in 13 papers. The themes of these papers are carried over as articles in this special section, either individually or in combination. The Journal of Aircraft editors, editorial staff, and reviewers are also thanked for the many helpful suggestions made during the publication process.

Polypyridylpropyne‐Pd and ‐Pt porphyrin coating for visualization of oxygen pressure

AbstractPoly[1‐trimethylsilyl‐1‐propyne‐co‐1‐(3‐pyridyl)propyne] 1 was prepared both as a polymer‐ligand of a palladium porphyrin (PdOEP) and of a platinum porphyrin (PtTFP) and as a highly gas‐permeable polymer matrix of the porphyrin. The porphyrin acted as a phosphorescence probe which could be quenched with oxygen and sense the oxygen partial pressure. 1 gave a smooth and tough coating with a thickness of ca. 2 µm which homogeneously involved the porphyrin. The porphyrin‐1 coatings displayed strong red‐colored phosphorescences (the emission maximum at 670 and 650 nm for PdOEP and PtTFP, respectively), and their intensity significantly decreased with an increase in the oxygen partial pressure on the coating. The high oxygen‐quenching efficiency or the high oxygen pressure sensitivity of the porphyrin's phosphorescence was observed even at cryogenic temperature. Aggregation of the porphyrin was suppressed in the coating by ligation of the porphyrin with the nitrogenous residue of 1 to significantly reduce spatial noise in the phosphorescence measurement or the oxygen‐pressure sensing. PtTFP‐1 was coated on the surface of a delta wing model. The oxygen‐pressure distribution on the coated model was successfully visualized in a cryogenic wind tunnel test. Copyright © 2008 John Wiley & Sons, Ltd.

1603 アペックスフェンスによるデルタ翼の空力特性の改善 : フェンス取り付け角の影響(G05-1 流体工学(1)物体・生体まわりの流れ,21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)

Apex fences are applied to improve low-speed aerodynamic characteristics of 60 deg. delta wing model. The longitudinal forces and moment acting on the thin delta wing with apex fences (Apex fences are used hinged to the delta wing upper surfaces along the leading edges at angles of 45,90 and 135 degrees.) are measured by using universal multi-component sensor and static pressure distribution on the upper surface is also measured by using pressure transducer with pressure scanner. Combination of total-pressure distribution maps and longitudinal force and moment measurements provide information on the three-dimensional nature of the vortex structure induced by apex fences.

앞전 형상에 따른 삼각 날개의 공력 특성

Flow visualization and aerodynamic characteristics of delta wings with two different leading edge geometries are investigated by PIV system and wind tunnel balance when the Reynolds number is about based on the freestream velocity and the root chord length. Delta wing models have 65-deg swept angle, and the leading edge shapes are divided into round- and sharp- type. The experimental results indicated that the leading-edge vortex strength and aerodynamic coefficient in the round leading edge are stronger and more, respectively than those in the sharp one. Therefore the flow interactions between vortices and the boundary layer are more desirable or more rapidly swirled in the round-type leading edge.

Pressure sensitive paint measurements on a delta-wing in supersonic flow

AbstractExperiments have been performed documenting the pressure field on the lee-side of a delta-wing at three incidence angles (5°, 10°, and 15°) and at Mach 1·8 using a PSP (Pressure Sensitive Paint) technique. The delta-wing model having a leading edge sweep of 60° was instrumented with 31 spanwise pressure ports at 68% of mean chord location. The Optrod-B1 binary paint was utilised and the PSP images were processed employing a resection based methodology. The comparisons of PSP results with those measured employing pressure taps show good agreement at different incidence angles.

Support Interference for a Maneuvering Delta Wing

The results of a recent investigation of support interference in dynamic wind-tunnel testing are presented. Particular emphasis has been focused on studying the interaction of vortices generated by delta wing models at high angle of attack with a support structure. A novel approach was used whereby the effects of time-varying vortex strength could be separated from the effects of the vortex-support interaction. With this approach, it was shown that the principal effect of support interference was to cause an upstream shift in the mean breakdown location, regardless of the form of motion or forcing frequency. The magnitude of the fluctuations in the breakdown location was shown to be dictated by support interference at low forcing frequencies, but by time-varying vortex strength effects at high frequencies. A number of important observations were also made of the vortex behavior at high forcing frequencies. In particular, a curious jumping behavior of the vortex breakdown location was observed. The presence of the support reduced the magnitude of the breakdown jump, but jumping was, nevertheless, observed in all of the cases considered at high enough forcing frequencies.

Modeling Delta Wing Limit-Cycle Oscillations Using a High-Fidelity Structural Model

Flutter and limit-cycle oscillations(LCO) of a delta-wing model are studied theoretically and correlated with results from an earlier experiment and an earlier simpler theoretical model. The present theoretical model uses a high-fidelity nonlinear structural model and a linear vortex lattice aerodynamic model. The commercial finite element package ANSYS is selected to model the structure and is coupled to the vortex lattice aerodynamic model using a subiteration procedure to carry out time simulations. The delta-wing model is studied for five angles of attack (0, 1, 2, 3, and 4 deg) and for various flow speeds. Theoretical results are calculated for two different root-chord boundary conditions, that is, fully fixed and also another that allows some in-plane movement at the root chord by attaching stiff in-plane springs that connect the structure to the root boundary. The results obtained using the high-fidelity structural model are compared to earlier results computed using a lower-fidelity von Karman plate theory. For all angles of attack studied here, the correlation between theory and experiment is better for the aeroelastic model, which uses the high-fidelity (ANSYS) structural model. Both flutter velocity and frequency as well as the LCO amplitudes and frequencies that are predicted using the higher-fidelity stuctural model correlate well with experiment. In particular the flutter and LCO results predicted using the high-fidelity structural model are similar, both qualitatively and quantitatively, for the two different in-plane boundary conditions. However the results obtained from the von Karman model differ substantially for the two different in-plane boundary conditions.

Microsensors and Actuators for Macrofluidic Control

Microsensors and actuators suitable for macrofluidic control have been designed, fabricated, tested, and optimized over the span of the last decade. MEMS-based shear stress sensor arrays using polysilicon hot filaments have been fabricated on both rigid (silicon) and flexible (parylene) substrate for application on all types of fluid dynamic and aerodynamic surfaces. In addition, MEMS bubble flap-type pneumatic actuators have been tested and used in turbulent boundary layer drag reduction in conjunction with the rigid MEMS shear stress sensor arrays acting as high-speed shear stress imagers. The flexible MEMS bubble actuator arrays have also been used with the flexible shear stress sensor arrays for generating maneuvering forces in the wind tunnel for a delta wing model and on UAV-type radio-controlled aircraft.

Temperature correction of PSP measurement using dual-luminophor coating

We developed a dual-luminophor pressure/temperature sensitive paint (DPTSP) to correct the temperature dependence of pressure-sensitive paint. The DPTSP is composed of two sensor molecules, PtTFPP and Rhodamine B (RhB), and Poly-IBM-co-TFEM as a binder. Temperature was determined from the image of RhB, and the temperature dependence of PtTFPP was corrected using the calculated temperature. To validate the capability of DPTSP for temperature correction, the pressure field on a delta wing model was measured by the DPTSP measurement system. The pressure values obtained with DPTSP were in good quantitative agreement with pressure tap data. It has been validated that DPTSP is effective in correcting the temperature dependence of PSP.