Double-delta Wing Research Articles

Dye visualization over double delta wing with various kink angles

In this study, the aim is to exhibit vortical behaviors of flow on double delta wings having 70° strake sweep angle and kink angles of 150°, 160° and 165° using dye visualization technique in a water channel. Experiments were performed at Reynolds numbers based on the chord length Re = 10,000 and 25,000 for angle of attack in the range 5° to 35°. The visualizations were performed for both end-view and cross-flow planes. The results revealed that the kink angle has a significant role on the interaction of vortices and the strake vortex breakdown locations. The interaction between the strake vortex and the wing vortex is dominant on the flow behaviors at α ≤ 10°. The flow behavior is affected by the kink angle. Two interaction mechanisms which are spiral and enveloping are observed. The spiral interaction alternates to enveloping interaction with increasing Reynolds number. Moreover, the trajectory of the strake vortex core moves outboard with increasing Reynolds number at α = 10°. For α ≥ 15°, Reynolds number is less effective on the strake vortex breakdown location and also the vortex breakdown locations move the apex gradually with increasing angle of attack. Wake-alike flow structure takes place after occurrence of the vortex breakdown since vortex core splits into disorganized small-scale vortices. On the other hand, development of the wing vortex is more complex than the strake vortex since it collapses near the vicinity of the junction.

Dynamic response of vortex breakdown flows to a pitching double-delta wing

A finite volume-based solver with rigid moving mesh and delayed detached eddy simulation (DDES) techniques is implemented to investigate the unsteady flows around an 80°/65° double-delta wing (DDW) subjected to sinusoidal pitching motions. The focus concentrates on understanding the behaviour of the burst point (BP), helical mode instability, pressure fluctuations, and dynamic pitching stability. The role of the reduced frequency (RF) in the response of the above features is analyzed in detail. It is in consistence with the previous experiments that the movement of the BP is nearly a simple harmonic motion and locked with the frequency of the pitching motion accompanied with a phase lag that is strongly affected by the RF. It is found that the time-averaged flow in the post-breakdown regions is approximately a conical flow, whose cone angle also depends on the RF. The natural frequency of BP oscillation at the stationary state of AOA = 36° is the critical frequency, which determines the sign of dynamic pitching derivative. A pair of critical frequencies is found, by which the features of the BP, cone angle of the helical structures, and dynamic pitching stability are divided into several linear sections. Two simplified 1st and 2nd order differential models are proposed and applied for predicting the dynamic behaviour of the BP, and the 2nd order model can give coincident hysteresis loops with the present computational fluid dynamics (CFD) results in a certain range.

Effect of leading edge shapes on 81°/45° double-delta wing at low speeds

Investigations were performed on an 81°/45° sweep double-delta wing at a freestream velocity of 20 m/s. Experiments consisted of the measurement of forces, static pressures, and surface flow visualizations. Effect of the leading edge shapes of the double-delta wing was studied. Results indicated a strong influence of the leading edge shape on the aerodynamic performance of the body. The increase in the bluntness of the leading edge augments the suction pressure and delays the vortex lift phenomena at higher angles of attack, which in turn enhances the lift over the wing. A reasonable agreement between the experiments and computations were observed.

연기선과 미세 수적을 이용한 두 가지 가시화 기법과 와류에의 적용

두 가지의 비표면 가시화 기법과 이들 기법의 와류 적용 예를 기술한다. 두 가지 가시화 기법 중 하나는 전통적인 연기선 기법이고, 또 다른 하나는 가정용 초음파 가습기에서 발생하는 미세 수적을 사용하는 기법이다. 연기선 기법은 공기 흐름 속도에 제한이 있으며(0.07 mm 지름의 연기선의 경우 약 5 m/sec), 오염의 문제점이 있지만 매우 정교하고 선명한 연기 유맥선 시트를 발생시킬 수 있다. 이 기법은 3차원 날개의 날개 끝 와류의 가시화에 적용되었다. 초음파 가습기 미세 수적 기법은 연기선 기법에 비하여 공기 흐름 속도를 보다 크게 할 수 있으며(10 m/sec 이상), 독성과 오염의 문제를 해결할 수 있다. 초음파 가습기 미세 수적 기법은 정점 스트레이크를 가지는 이중 삼각날개에서 발생하는 복잡한 와류 시스템의 가시화에 성공적으로 적용되었다. The present paper describes the two off-surface visualization techniques and their application examples to vortex flows. One of the two visualization techniques is the classical smoke-wire technique, and the other is the visualization technique using the micro water-droplets generated by the home-style ultrasonic humidifier. The smoke-wire technique has the limit of air flow speed (about 5 m/sec for 0.07 mm-diameter wire) and the pollution problem, but it produces very fine and clear streak line sheet. It is applied to visualize the wing-tip vortices of a 3-dimensional wing. The micro water-droplet technique has the larger limit of air flow speed (above 10 m/sec) and is free from pollution and toxic problems compared to the smoke-wire technique. It is successfully applied to visualize the complex vortex system of a double-delta wing with an apex strake.

EFFECT OF GROUND PROXIMITY ON AERODYNAMIC FORCES AND MOMENTS OF REUSABLE LAUNCH VEHICLE AT SUBSONIC SPEED

This paper explores the effect of ground proximity on aerodynamics forces and moments acting on reusable launch vehicle at the time of its landing. As the name indicates the concept of reusability is attractive especially in case of Space Transportation and launch vehicles as it reduces the mission cost. The design of Reusable Launch vehicle is also very critical as the vehicle has to pass through various speed ranges from hypersonic to subsonic. RLV wing is generally low aspect ratio double delta wing which is very sensitive to ground effect due to the characteristic excessive vehicle attitudes associated with the landing and touchdown phases of flight.. In present study, a RLV model of 260 mm total length and reference area of 155mm2 had been modeled using CAD tools. It has double delta wing of wing span 144mm and mean chord of 76mm and vertical stabilizer of height 38mm. Effect of ground proximity on aerodynamic forces acting on RLV has been studied using commercial CFD software. During present analysis computational analysis has been done for different height. Height of RLV is varied from free flight condition to landing touchdown height and the angle of attack has been varied for 0 to 20. The complete study is done for lower subsonic speed at Reynolds number of 1.8*106 for fix ground and moving ground conditions separately.

Experimental study of flow field distribution over a generic cranked double delta wing

The flow fields over a generic cranked double delta wing were investigated. Pressure and velocity distributions were obtained using a Pitot tube and a hot wire anemometer. Two different leading edge shapes, namely “sharp” and “round”, were applied to the wing. The wing had two sweep angles of 55° and 30°. The experiments were conducted in a closed circuit wind tunnel at velocity 20m/s and angles of attack of 5°–20° with the step of 5°. The Reynolds number of the model was about 2×105 according to the root chord. A dual vortex structure was formed above the wing surface. A pressure drop occurred at the vortex core and the root mean square of the measured velocity increased at the core of the vortices, reflecting the instability of the flow in that region. The magnitude of power spectral density increased strongly in spanwise direction and had the maximum value at the vortex core. By increasing the angle of attack, the pressure drop increased and the vortices became wider; the vortices moved inboard along the wing, and away from the surface; the flow separation was initiated from the outer portion of the wing and developed to its inner part. The vortices of the wing of the sharp leading edge were stronger than those of the round one.

Interaction of multiple vortices over a double delta wing

Interaction of strake and wing vortices over a 70°/50° double delta wing were studied experimentally in a wind tunnel using particle image velocimetry (PIV) measurements. The upstream effect of the wing vortex on the formation of the strake vortex was identified. A dual-vortex structure of the strake vortices was observed before the wing vortex developed. Further downstream, wing and strake vortices rotated around each other slowly initially, and then faster with downstream distance, at an increasing rate with increasing incidence. Prior to vortex breakdown, both wing and strake vortices were found meandering in relatively small regions. The correlation between the instantaneous locations of the vortices increases if the vortices become sufficiently close to each other. The proper orthogonal decomposition (POD) analysis of the instantaneous velocity fields suggested that, for both wing and strake vortices, the most energetic mode was displacement in the first helical mode. The most energetic mode reveals out-of-phase displacements when the vortices are close to each other.

Effects of pitch, yaw, and roll on delta wing skin friction topology

High-resolution skin friction fields in separated flows over a 65° delta wing and a 76/40° double-delta wing with different junction fillets are obtained by using quantitative global skin friction diagnostics based on luminescent oil visualizations. The topological features such as separation and attachment lines on these delta wings are identified, and the effects of the pitch, yaw, and roll angles on the skin friction topology are studied systematically. The topological consistency of the measured skin friction fields on the delta wings is examined by using the Poincare–Bendixson index formula.

Skin Friction Fields and Surface Dye Patterns on Delta Wings in Water Flows

This paper discusses the relationship between skin friction fields and surface dye patterns in surface luminescent dye visualizations in water flows, providing a theoretical foundation for extraction of high-resolution skin friction fields. The limiting form of the mass diffusion equation at a wall is recast as an optical flow equation connecting skin friction with the luminescent dye intensity. Snapshot solutions are obtained from a time sequence of luminescent intensity images by solving the optical flow equation via the variational method, and then a normalized skin friction field is reconstructed by averaging the snapshot solutions. An error analysis is given to identify the major error sources and the limitations of the technique. To evaluate the feasibility of this technique, surface luminescent dye visualizations on a 65 deg delta wing and a 76/40 deg double-delta wing are conducted in a water tunnel. The extracted skin friction topology on the delta wings and the velocity fields obtained by using particle image velocimetry (PIV) are discussed.

Numerical Simulation of 650 Delta Wing and 650/400 Double Delta Wing to Study the Behaviour of Primary Vortices on Aerodynamic Characteristics

Numerical Simulation of 650 Delta Wing and 650/400 Double Delta Wing to Study the Behaviour of Primary Vortices on Aerodynamic Characteristics

Challenges in the development of a slow burning solid rocket booster

Solid rocket motors always find a place in launch vehicles, especially as booster stages on account of their capability to deliver higher thrust needed during the initial phase of flight, most importantly during lift-off phase. The newly developed slow burning solid booster is chosen for the first stage of Reusable Launch Vehicle-Technology Demonstration (RLV-TD) flight programme. This suborbital two stage technology demonstration programme is aimed for a hypersonic re-entry of double delta wing vehicle from an altitude of nearly 65–70 km, where the first stage of vehicle is identified as solid booster. Therefore, the desirable Mach number of the vehicle at motor burnout (30–35 km) is essentially required to be minimum 6 to attain hypersonic re-entry. The contradictory mission requirements like achieving the desired Mach number and limiting the dynamic pressure demanded for design, development and qualification of a new motor with slowest burning solid propellant thereby an action time of nearly 90 s. The challenges faced in the design, development, realization and qualification of the motor through two successful static tests are discussed in this paper. The special attention given for motor case with 2 mm shell thickness, development of a slow burn rate propellant, ignition strategies for the slow burn rate propellant, stable burning of the low burn rate propellant and flow separation related issues for ground test are highlighted.

Numerical investigation of unsteady vortex breakdown past 80°/65° double-delta wing

An improved delayed detached eddy simulation (IDDES) method based on the k-ω-SST (shear stress transport) turbulence model was applied to predict the unsteady vortex breakdown past an 80°/65° double-delta wing (DDW), where the angles of attack (AOAs) range from 30° to 40°. Firstly, the IDDES model and the relative numerical methods were validated by simulating the massively separated flow around an NACA0021 straight wing at the AOA of 60°. The fluctuation properties of the lift and pressure coefficients were analyzed and compared with the available measurements. For the DDW case, the computations were compared with such measurements as the mean lift, drag, pitching moment, pressure coefficients and breakdown locations. Furthermore, the unsteady properties were investigated in detail, such as the frequencies of force and moments, pressure fluctuation on the upper surface, typical vortex breakdown patterns at three moments, and the distributions of kinetic turbulence energy at a stream wise section. Two dominated modes are observed, in which their Strouhal numbers are 1.0 at the AOAs of 30°, 32° and 34° and 0.7 at the AOAs of 36°, 38° and 40°. The breakdown vortex always moves upstream and downstream and its types change alternatively. Furthermore, the vortex can be identified as breakdown or not through the mean pressure, root mean square of pressure, or even through correlation analysis.

Numerical Investigations for the Effect of Slender Body on Dynamic Rolling Characteristics of a 80°/60° Double Delta Wing

Numerical Investigations for the Effect of Slender Body on Dynamic Rolling Characteristics of a 80°/60° Double Delta Wing

Numerical Investigations for the Effect of Slender Body on Dynamic Rolling Characteristics of a 80°/60° Double Delta Wing

The effect of slender body on the rolling characteristics of a double delta wing is found by comparing the numerical simulation results of the double delta wing and wing-body configuration. The coupled computation system solving the Navier-Stokes equations and the rolling motion equation alternatively to obtain the unsteady vortical flow around the two configurations while rolling. The results conclusively showed the upwash effect of the slender body enhanced the energy of strake vortex and merged vortex.The aerodynamic lag of double delta wing is weak, contrarily, the time lag effect of the wing-body configuration is significant. The asymmetry vortices structure nearby the trailing edge are believed to be the main reason for the unsteady time lag effect.

Detached-eddy simulations on massively separated flows over a 76/40° double-delta wing

We investigated the flow past a 76/40° double-delta wing at angles of attack from 10° to 35° using SA-DES and a hybrid fourth-order scheme. Comparisons were made with the surface separation pattern, the pressure distribution and the location of vortex breakdown with experimental data, and acceptable agreement was obtained. Further discussions were made under the scenario of massively separated flows on topics such as the characteristics of conical flow, the structure of vortex breakdown and vortical substructures.

Detached-eddy simulations and analyses on new vortical flows over a 76/40° double delta wing

Numerical simulations were performed on the massively separated flows of a 76/40° double delta wing using detached-eddy simulation (DES). A new type of cross-flow vortex is suggested. A vortex was initially generated near the junction of the strake and wing, which then moved towards the wing tip at certain wavelength and speed. Analyses were made in detail on the mechanism of the generation of the cross-flow vortex, that is, the inviscid cross-flow instability which differs from that of the swept blunt wing. Cross-section topology of the cross-flow vortex is also investigated, and the wavelength of the vortex array and the characteristic frequency are given. The analyses showed that the cross-flow vortices have an influence on the pressure distribution, which can cause a 10%–20% deviation from the averaged distribution.

Supersonic Testing of NAL Binary Pressure Sensitive Paint

The main objective of the present study is to subject NALPSP-02, an indigenously developed binary pressure sensitive paint, to supersonic flow conditions. An experimental study was carried out on a double-delta wing model at a free stream Mach number of 2.0 in the 0.3 m trisonic wind tunnel at NAL. Pressure Sensitive Paint (PSP) technique was used to measure the pressure distribution on the lee side of the double delta wing at an incidences of 0°, 5°, 10° and 15°. The PSP results were compared against surface pressure data obtained using conventional pressure taps on the same model.

Visualisation of Vortex Structures Developed on the Upper Surface of Double-Delta Wings

A large number of studies of flow visualisations, developed on the upper surface of delta and of gothic wings, and of cones, have been carried out in the wind tunnel of the Valenciennes University’s Aerodynamics and Hydrodynamics Laboratory. These studies have provided a better understanding of the development and of the positioning of vortex structures and have enabled, in particular, the preferential nature of intervortex angles, thereby defined, to be determined. This study of the vortical structures developed on the upper surface of a double-delta wing has revealed that these vortex flows are quite complex and that vortex structures interact with one another. Indeed, it would seem that vortex behaviour has something of a universal nature. An angular conformity between primary and secondary vortex torques and the leading edges of the wing can be expressed by the law of filiation. Intervortex angles evolve with increasing incidence while fragmentation is a function of the apex angle. It would be interesting to recall that this particular spatial organisation of vortex structures, citing the concept of preferential angles, also appears in standard theories on aerodynamics as, for example, in those governing aerodynamic drag. Nevertheless, the link between interior and exterior vortex structures remains to be investigated further. Such studies might even prove the existence of a supplementary torque. In addition, the least resistance of secondary vortices in relation to their fragmentation inevitably calls for experiments to be undertaken with other possible combinations of slender bodies although these areas of research are beyond the scope of this article.

Special Section on Intelligent Unmanned Systems

There has been increased worldwide interest in the research on intelligent unmanned systems in recent decades. The unmannedsystem technologies have been implemented in the underwater, ground, aerial, and space domains within an ever-extending operational envelope. The miniaturization of the hardware footprint coupled with exponentially increased computing power has enabled more efficient application of a wide range of unmanned vehicles. The papers in this special section represent recent contributions to several aspects of intelligent unmanned systems. The papers were largely an extension of select papers presented at the ICIUS2010 held in Bali, Indonesia, in October 2010. Three papers deal with the development of intelligent algorithms for unmanned aerial vehicles, while two papers focus on the software and hardware integration. The rest of the papers cover the application of aerospace technology for unmanned-vehicle design. The paper by Manathara and Ghose addresses the problem of multiple unmanned aerial vehicle (UAV) rendezvous when the UAVs have to perform maneuvers to avoid collisions with other UAVs. The authors proposed a solution that consists of velocity control and a wandering maneuver of the UAVs based on a consensus among them on the estimated time of arrival at the point of rendezvous. This simple algorithm was shown to be useful in tracking stationary or slowly moving targets with a standoff distance. Kim et al. paper present an automatic landing control design using adaptive, integrated guidance and control (IGC) logic. The proposed IGC design uses a combination of an adaptive output feedback inversion and backstepping techniques. The authors formulated the problem as an adaptive output feedback control problem for a line-of-sight–based chasing-flight configuration. The automatic landing system using IGC logic was evaluated using a sophisticated six-degrees-of-freedom nonlinear simulation program with the approach and landing scenario. Jeong et al. focused on the design of a robust H-infinity attitude controller for an unmanned small-scale helicopter. It was demonstrated that the H-infinity controller works amiably when applied to the nonlinear model, even though it is designed using a linear model approximation. The time-domain simulation with the nonlinear model demonstrated that the proposed controller was very robust to the uncertainties as expected, overcoming large gain uncertainties and time delay in each input channel. Formal verification currently plays an important role in guaranteeing the correctness of the operation of safety-critical systems, for example, in UAVs. Lee et al. discuss formal modeling and verification of operational flight programs (OFPs) in a small-scale unmanned helicopter. Two different formal models for the OFP were developed and verified using SPIN and UPPAAL model checkers. Real-time behavior of theOFPwas specified using a timed-automata model. Their verification of the OFP found several safety-critical faults to be reported to and fixed by the development team. One of the key features of autonomous flight of UAVs is the ability to perform localization and map building of flight environments simultaneously. Park et al. present vision-based simultaneous localization and mapping (SLAM) for small UAVs in global-positioningsystem–denied environments. A method that combines onboard and on-ground localization and mapping was proposed. The paper by Thu et al. investigates the nature of vortex structure developed by a double-delta wing. Dye-flow visualization in a small-scale water tunnel was used for the study. The authors identified two types of flow interaction, which are enveloping and coiling-up modes. Characteristics of the two modes were examined for various angles of attack (AOAs) at relatively low Reynolds numbers of 10,000 and 15,000. The authors found that the wing vortex became stronger and the enveloping mode was dominant for a low AOA, while the coiling-up mode was dominant at a high AOA. They also reported that the transition of these two modes occurred at around an AOA of 20 . Another paper in the special section studied the unsteady flow field of a flapping model butterfly under the zero-advance ratio (J 5 0) condition using timeresolved stereoscopic particle image velocimetry. The effect of feathering on thrust generation because of chordwise wing flexibility was identified by focusing on the ejection of airflow from the gap between two wings. Ghosh et al. found that the wing flexibility adds a feathering effect and enhances net lift and thrust generation through ejection of vortex pairs during the downstroke-to-upstroke–motion reversal. They also reported that the thrust produced was positive all over the cycle, even at J 5 0, because of the feathering motion. The special section editors would like to express their sincerest gratitude to all the contributing authors. We hope that the readers will find the content of this section not only relevant but also stimulating and beneficial for their ongoing research endeavors.

High-Order and High Accurate CFD Methods and Their Applications for Complex Grid Problems

AbstractThe purpose of this article is to summarize our recent progress in high-order and high accurate CFD methods for flow problems with complex grids as well as to discuss the engineering prospects in using these methods. Despite the rapid development of high-order algorithms in CFD, the applications of high-order and high accurate methods on complex configurations are still limited. One of the main reasons which hinder the widely applications of these methods is the complexity of grids. Many aspects which can be neglected for low-order schemes must be treated carefully for high-order ones when the configurations are complex. In order to implement high-order finite difference schemes on complex multi-block grids, the geometric conservation law and block-interface conditions are discussed. A conservative metric method is applied to calculate the grid derivatives, and a characteristic-based interface condition is employed to fulfil high-order multi-block computing. The fifth-order WCNS-E-5 proposed by Deng is applied to simulate flows with complex grids, including a double-delta wing, a transonic airplane configuration, and a hypersonic X-38 configuration. The results in this paper and the references show pleasant prospects in engineering-oriented applications of high-order schemes.