Blunt Leading Edges Research Articles

Effects of high-temperature and entropy layer of Mach 12 oblique shock wave

In the current study, we numerically investigate the high-temperature inviscid flows over the 15-degree wedges with the sharp and blunt leading-edges. Firstly, the vibrational non-equilibrium effects in sharp wedge flow with different deflection angles are analyzed. Then, the influence of the vibrational non-equilibrium effects on the flow parameters behind the oblique shock wave is provided. Lastly, we discuss the high-temperature non-equilibrium flow over the 15-degree wedge with the entropy layer.

Numerical Simulation of the Interaction between Weak Shock Waves and Supersonic Boundary Layer on a Flat Plate with the Blunt Leading Edge

The interaction of weak shock waves in form of an N-wave with the supersonic laminar boundary layer on a flat plate with blunt leading edge at the free-stream Mach number M = 2.5 is numerically studied. The numerical results are compared with known experimental data. The combined influence of the N-wave and the leading edge bluntness on the laminar-turbulent transition process is discussed.

Shock Reduction through Opposing Jets—Aerodynamic Performance and Flight Stability Perspectives

In this research paper, investigations of counter flow (opposing) jet on the aerodynamic performance, and flight stability characteristics of an airfoil with blunt leading-edge in supersonic regime are performed. Unsteady Reynolds-Averaged Navier-Stokes ( U R A N S ) based solver is used to model the flow field. The effect of angle of attack ( α ), free-stream Mach number ( M ∞ ), and pressure ratio ( P R ) on aerodynamic performance of airfoil with and without jet are compared. The results indicate that the opposing jet reduces drag from 30 % to 70 % , improves the maximum lift-to-drag ratio from 2.5 to 4.0, and increases shock stand-off distance from 15 % to 35 % depending on flow conditions. The effect of opposing jet on longitudinal flight stability characteristics, studied for the first time, indicate improvement in dynamic stability coefficients ( C m q + C m α ˙ ) at low angles of attack. It is concluded that the opposing jet can help mitigate flight disturbances in supersonic regime.

Leading-Edge Roughness Affecting Diamond-Wing Aerodynamic Characteristics

Diamond wing configurations for low signature vehicles have been studied in recent years. Yet, despite numerous research on highly swept, sharp edged wings, little research on aerodynamics of semi-slender wings with blunt leading-edges exists. This paper reports on the stall characteristics of the AVT-183 diamond wing configuration with variation of leading-edge roughness size and Reynolds number. Wind tunnel testing applying force and surface pressure measurements are conducted and the results presented and analysed. For the investigated Reynolds number range of 2.1 × 10 6 ≤ R e ≤ 2.7 × 10 6 there is no significant influence on the aerodynamic coefficients. However, leading-edge roughness height influences the vortex separation location. Trip dots produced the most downstream located vortex separation onset. Increasing the roughness size shifts the separation onset upstream. Prior to stall, global aerodynamic coefficients are little influenced by leading-edge roughness. In contrast, maximum lift and maximum angle of attack is reduced with increasing disturbance height. Surface pressure fluctuations show dominant broadband frequency peaks, distinctive for moderate sweep vortex breakdown. The experimental work presented here provides insights into the aerodynamic characteristics of diamond wings in a wide parameter space including a relevant angle of attack range up to post-stall.

The Effect of Aerodynamic Evaluators on the Multi-Objective Optimization of Flatback Airfoils

With the long lengths of today's wind turbine rotor blades, there is a need to reduce the mass, thereby requiring stiffer airfoils, while maintaining the aerodynamic efficiency of the airfoils, particularly in the inboard region of the blade where structural demands are highest. Using a genetic algorithm, the multi-objective aero-structural optimization of 30% thick flatback airfoils was systematically performed for a variety of aerodynamic evaluators such as lift-to-drag ratio (Cl/Cd), torque (Ct), and torque-to-thrust ratio (Ct/Cn) to determine their influence on airfoil shape and performance. The airfoil optimized for Ct possessed a 4.8% thick trailing-edge, and a rather blunt leading-edge region which creates high levels of lift and correspondingly, drag. It's ability to maintain similar levels of lift and drag under forced transition conditions proved it's insensitivity to roughness. The airfoil optimized for Cl/Cd displayed relatively poor insensitivity to roughness due to the rather aft-located free transition points. The Ct/Cn optimized airfoil was found to have a very similar shape to that of the Cl/Cd airfoil, with a slightly more blunt leading-edge which aided in providing higher levels of lift and moderate insensitivity to roughness. The influence of the chosen aerodynamic evaluator under the specified conditions and constraints in the optimization of wind turbine airfoils is shown to have a direct impact on the airfoil shape and performance.

The Effect of an Eroded Leading Edge on the Aerodynamic Performance of a Transonic Fan Blade Cascade

Especially at transonic flow conditions the leading edge shape influences the performance of a fan profile. At the same time the leading edge of a fan profile is highly affected by erosion during operation. This erosion leads to a deformation of the leading edge shape and a reduction of the chord length. In the present experimental and numerical study, the aerodynamic performance of an original fan profile geometry is compared to an eroded fan profile with a blunt leading edge (BLE) and a chord length reduced by about 1%. The experiments are performed at a linear fan blade cascade in the Transonic Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is 1.25 and the Reynolds number 1.5 × 106. All tests are carried out at a low inflow turbulence level of 0.8%. The results of the investigation show that losses are increased over the whole operating range of the cascade. At the aerodynamic design point (ADP) the losses raise by 25%. This significant loss increase can be traced back to the increase of the shock losses at the leading edge. The change in shock structure is investigated and described in detail by means of particle image velocimetry (PIV) measurements and Schlieren imaging. Additionally, the unsteady fluctuation of the shock position is measured by a high-speed shadowgraphy. Then the frequency range of the fluctuation is obtained by a Fourier analysis of the time resolved shock position. Furthermore, liquid crystal measurements are performed in order to analyze the influence of the leading edge shape on the development of the suction side boundary layer. The results show that for the original fan blade the transition occurs at the shock position on the blade suction side by a separation bubble whereas the transition onset is shifted upstream for the fan blade with the BLE.

Open-Type Separation on Delta Wings for Leading-Edge Bluntness

A numerical simulation was carried out corresponding to recent experiments using delta wings with sharp and blunt leading-edges, which indicates the second primary vortex, at NASA Langley Research Center. However, the experimental data did not reveal the detailed physical phenomena regarding the second primary vortex, because the experiment used only on-the-body-surface data. In the present study, the physical phenomena were revealed using Reynolds-averaged Navier-Stokes computations with three one-equation turbulence models on an unstructured hybrid mesh. The adaptive mesh refinement method in the vicinity of the vortex center was also applied to have more mesh resolution. Consequently, the result quantitatively revealed that appropriate modeling regarding turbulent kinematic viscosity was significant. Moreover, the three-dimensional visualization of the computational fluid dynamics results suggested that the second primary vortex was a developing shear layer merging into an open-type separation generated late by the primary vortex.

The effect of inlet conditions on flow over backward facing step

The objectives of this paper are to present measurements of physical phenomena associated with two-dimensional separation of a turbulent boundary layer over a backward-facing step (B.W.F.S.). Two flow conditions were studied: Sharp leading edge (S.E.) with flow underneath (Case I) and blunt leading edge (B.E.) with flow underneath (Case II). The comparisons between the two cases are discussed with reference to changes in: (i) The potential flow. (ii) The free shear layer surrounding the bubbles. (iii) The separation condition (separation point, shape of bubbles, reattachment point, etc.). All the experiments on the full scale model were carried out inside an open-circuit low-speed wind tunnel. The experiments were carried out at a Reynolds number of Re H =7.9×10 4 based on step height H . Measurements were made of the longitudinal, U ̄ , and vertical, V ̄ , components of the mean velocity, the two normal stress components u ̄ 2 and, v ̄ 2 and the shear stress uv . These data were obtained using the Flying Hot-Wire (F.H.W.) anemometer. The F.H.W. anemometer can be used with minimal flow interference. The experimental investigations have shown significant differences between the recirculating regions associated with the two conditions studied due to different inlet conditions.

Experimental Results on a Mach 14 Waverider with Blunt Leading Edges

The results of wind-tunnel tests of a Mach 14 waverider were analyzed to assess its overall performance. The waverider was optimized using a e gure of merit that included viscosity, volume, and lift-to-drag considerations. The e nal design included a 0.25-in. leading-edge radius. The general performance of the tested waverider was analyzed and compared to the theoretical design from which it was derived. The theoretical design was generated from a conical e owe eld, where the vehicle’ s ine nitely sharp leading edges were everywhere-attached to the conical shock. Since the tested waverider featured blunt leading edges, it was important to assess the performance losses associated with e ow spillage; these losses were found to be relatively small. However, the increased drag due to the blunt leading edges contributed greatly to reducing the aerodynamic performance of the tested waverider. Also, it was found that the aerodynamic coefe cient data were insensitive to changes in Mach number and Reynolds number, indicating excellent off-design performance for the ranges of values tested.

A New Technique for Analysis of Unsteady Aerodynamic Responses of Cascade Airfoils with Blunt Leading Edge. Unsteady Aerodynamic Responses of the Cascade in Incompressible Flow.

Unsteady aerodynamic responses of cascade airfoils with a blunt leading edge, which are subjected to incident rotational and irrotational fluctuations, are analyzed by use of the method developed in the previous paper of Funazaki and Kakudate. Since some difficulties in the imposition of the boundary condition have been pointed out in that paper, most of the efforts in this study are devoted to refinement of the method of the boundary condition imposition on the inlet/outlet and periodic boundaries by employing the analytical solutions in the upstream/downstream far-fields. Although the method presented in this paper could be applied to compressible flow problems, numerical examples employed this time are limited to incompressible flow cases. Obtained results then show the usefulness of the method, and also reveal some of the notable features of the unsteady flow field induced within the blade-to-blade passage vividly.

A New Technique for Analysis of Unsteady Aerodynamic Responses of Cascade Airfoils with Blunt Leading Edge. 1st Report. Theory.

A new technique has been developed, based on the technique proposed by Goldstein (1978), in order to analyze unsteady aerodynamic responses of cascade airfoils subjected to incident wakes. One of the important features of this technique is the introduction of the additional perturbation potential for the purpose of preventing the solution being indeterminate due to the singular behavior of the unsteady perturbed velocities around the blunt leading edge. In this paper, the main focus is placed on the theoretical development of the technique, and some simple cases of numerical calculations with no stagger angle are made in order to show the validity of our method.

Experimental Evaluation of the Effects of a Blunt Leading Edge on the Performance of a Transonic Rotor

A rotor designed for a tip relative Mach number of 1.4 and a pressure ratio of 1.8 was tested with two different leading edge configurations, to evaluate experimentally the effects of leading edge thickness on the performance of a transonic rotor. The rotor was first tested with a leading edge thickness that resulted in a normal blade gap blockage of approximately 3 percent at the rotor tip and increased to about 4 percent at the rotor hub. The rotor leading edge was then cut back to produce a leading edge thickness which resulted in a normal blade gap blockage of about 6 percent at the tip and the hub remained at about 4 percent. The increased leading edge thickness resulted in a decrease in the rotor overall peak efficiency of 3.5 points at design speed. The major portion of this loss in efficiency occurred in the outer 30 percent of the blade span. At 90 and 70 percent of design speed, the increased leading edge thickness had a relatively small effect on rotor efficiency.

Errata: "Numerical Method for Hypersonic Internal Flow over Blunt Leading Edges and Two Blunt Bodies"

Errata: "Numerical Method for Hypersonic Internal Flow over Blunt Leading Edges and Two Blunt Bodies"

Numerical Method for Hypersonic Internal Flow over Blunt Leading Edges and Two Blunt Bodies

The steady-state solution of mixed subsonic-supersonic flow near a blunt leading edge in hypersonic internal flow and supersonic flow over two blunt bodies is obtained as an asymptotic solution of the unsteady equations of motion by a time-dependent method. The continuous flow region between the shock and the body is calculated by the method of finite differences and exact boundary conditions are applied using the two-dimensional, unsteady, method of characteristics. Numerical results are presented for two-dimensional and axisymmetric flows with a continuously curved shock and shock interacting at the axis of symmetry as Mach or regular interaction. The shock wave configuration is found to be dependent on freestreain Mach number, leading edge shape and bluntness. Theoretical calculations for inviscid, perfect gas are compared with experimental data, obtained in a hypersonic gun tunnel at a freestream Mach number of 8.33, for leading edge shapes of circular cross section.

The Effects of Blunt Leading Edges on Delta Wings at Mach 5.8

Pressure distributions were measured on a series of four delta wings with subsonic and supersonic leading edges, both sharp and blunt. The blunt leading edge radius was about 0.5 per cent of root chord. Schlieren studies were also made to determine top and side view shock locations. The tests were conducted at a nominal Mach number of 5.8, and at Reynolds numbers between 0.335 x 10^6 and 0.901 x 10^6 based on root chord. Angular settings covered a range -0.2 ≤ w/V ≤ 0.5 in pitch at zero yaw {about -11.5° ≤ a ≤ + 30°), and a range of v/V = ± 0.125 (about ± 7.2°) at a fixed angle of pitch of 11.5°. The effects of bluntness were found to be small. Also, the pressures produced by shock wave interactions with the boundary layer, and the inviscid pressures generated by the blunt leading edges, were found to be small compared with the inviscid pressures producing lift on the basic wing. Spanwise pressure distributions show no similarity to those obtained by linearized theory. Centerline lower surface pressure in pitch at zero yaw is bracketed between the Newtonian value ΔP/q = 2(w/V)^2 and the two-dimensional exact value.

Inviscid Hypersonic Flow Over Blunt-Nosed Slender Bodies

At hypersonic speeds the drag/area of a blunt nose is much larger than the drag/area of a slender afterbody, and the energy contained in the flow field in a plane at right angles to the flight direction is nearly constant over a downstream distance many times greater than the characteristic nose dimension. The transverse flow field exhibits certain similarity properties directly analogous to the flow similarity behind an intense blast wave found by G. I. Taylor and S. C. Lin. Conditions for constant energy show that the shape of the bow shock wave R(x) not too close to the nose is given by R/d = K_1 (γ)(d/c)^(1/2) for a body of revolution, and by R/d = K_0(γ) (x/d)^(2/3) for a planar body, where d is nose diameter, or leading-edge thickness. A comparison with the experiments of Hammitt, Vas, and Bogdonoff on a flat plate with a blunt leading-edge at M_∞ = 13 in helium shows that the shock wave shape is predicted very accurately by this analysis. The predicted surface pressure distribution is somewhat less satisfactory. Energy considerations combined with a detailed study of the equations of motion show that flow similarity is also possible for a class of bodies of the form r_b ~ x^m, provided that m' ≤ m ≤ 1, where m' = 3/4 for a planar body and m' = (3/2(γ+1))/(3γ + 2) for a body of revolution. When m < m' the shock shape is not similar to the body shape, and except for the constant energy flows the entire flow field some distance from the nose must depend to some extent on the details of the nose geometry. Be again again utilizing energy and drag considerations one finds that at hypersonic speeds the inviscid surface pressures generated by a blunt nose are larger than the pressures produced by boundary layer growth on a flat surface over a distance from the nose of order l, where l/d ≃ 1/15 ((Re_d)/M_∞^2))^3 (Here Re_d is free-stream Reynolds number based on leading-edge thickness.) Thus at M_∞ = 15 the viscous interaction effects should be important for Re_d 3000 the inviscid pressure field is dominant and determines the boundary layer development, skin friction and heat transfer over the forward portion of the body. These rough estimates are in qualitative agreement with the experimental results of References 7 and 9.