Percent Of Chord Research Articles

EFFECT OF VARIABLE CANT ANGLE ON SWEEPBACK WING

Winglets are commonly used drag-reduction and fuel-saving technologies in today’s aviation. The primary purpose of the winglets is to reduce the lift-induced drag, therefore improving fuel efficiency and aircraft performance. Traditional winglets are designed as fixed devices attached at the tips of the wings. However, because they are fixed surfaces, they give their best lift-induced drag reduction at a single design point. In this work, we propose the use of variable cant angle winglets which could potentially allow aircraft to get the best all-around performance (in terms of lift-induced drag reduction), at different angle-of-attack values.This paper describes a XFLR 3-dimensional winglets analysis that was performed on a wing of NACA2412 cross sectional airfoil. The wing has span 2.40 m, root chord 0.180 m, tip chord 0.110 m, sweep angle 5 degree and taper ratio 0.556 and for a winglet NACA 0024 cross-sectional airfoil was considered of max thickness 24 percent at 30 percent chord, max camber 0 percent at 0 percent chord. The present study shows effects of wing without winglet, wing with winglet at cant angle 30, 60, 90.The results obtained from the analysis demonstrate that by carefully adjusting the cant angle, the aerodynamic performance can be improved.

Experimental Investigation of Spoiler Deployment on Wing Stall

Upper surface wing flaps, known as spoiler, are typically used to reduce lift and increase drag at touchdown; however spoilers have been shown to increase lift and reduce drag at near-stall conditions. The purpose of this experiment was to determine the spoilers’ impact on lift, drag, moment, and aerodynamic efficiency of a NACA 2412 airfoil at angles of attack (α) from −8 ° to 32 °. The experiment was conducted in the Ryerson Low-Speed Wind Tunnel (closed-circuit, 1 m × 1 m test section) at Re=783761, Ma=0.136. The lift coefficient (Cl), drag coefficient (Cd), moment coefficient about the quarter-chord () were captured with a changing spoiler deflection angle (δ) and spoiler length (b in percent chord). It was found that deflecting the spoiler resulted in an increase maximum lift of up to 2.497%. It was found that deflecting the spoiler by 8° was optimal for the b=10 cases. Any larger deflection reduced the lift gain, and a deflection of 25° caused the maximum lift to be 2.786% less than the clean configuration. In the b=15 case, δ=15° was optimal (1.760% maximum lift coefficient increase). The b=10 cases increased maximum lift coefficient between 0.35% and 2.10% higher than the b=15 cases. The source of the lift gain at high angles of attack is apparent in an analysis of the airfoil pressure distribution. The spoiler increased the suction peak on the airfoil surface upstream of the spoiler, and increased the pressure downstream. However the suction increase upstream is larger than the pressure increase downstream, resulting in a net increase in lift. The spoiler increased the stall angle 37.658% to 87.658% higher than the clean configuration. Stall angle increased with both δ and with an increased spoiler length. The spoiler airfoil produced less drag than the clean configuration at high angles of attack. The combination of the increased lift, and reduced drag resulted in an increase in aerodynamic efficiency at high angle of attack.

Vortical flow prediction of a diamond wing with rounded leading edges

The objective of this work is to assess the potential and limitations of current practice in computational fluid dynamic modeling in predicting vortical flowfields over a generic 53-degree swept diamond wing with rounded leading edges. This wing was designed under STO AVT Task Group 183 and has a constant NACA 64A-006 airfoil section with a leading edge radius of 0.264 percent chord. CFD simulations were run for different angles of attack at a Mach number of 0.15 and a Reynolds number of 2.7×106 based on the mean aerodynamic chord to match experiments. The wind tunnel experiments of the diamond wing were carried out in the Institute of Aerodynamics and Fluid Mechanics of the Technische Universität München, Germany and include aerodynamic lift, drag, and pitch moment measurements as well as span-wise pressure distributions at different chord-wise locations. This data set is used to validate the CFD results. The results presented demonstrate that the CFD compare well with the experiments at small angles of attack; the pitch moments predicted by the SARC turbulence model provide a better match to experimental results than the SA model at moderate angles of attack; and at high angles of attack, CFD predictions are not as good. The flow visualization results show that a leading-edge vortex is formed above the upper surface of the wing at an angle of attack of about eight degrees. This vortex becomes larger and stronger when the angle of attack is increased. With increasing angle of attack, the vortex formation point moves upstream and the vortex core moves inboard towards the wing center. Finally, the computational results show that the flow over the diamond wing is relatively steady throughout the range of angles of attack.

Impact of Nonuniform Leading Edge Coatings on the Aerodynamic Performance of Compressor Airfoils

A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and nonuniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge nonuniformity were added. A nonuniform coating is obtained when a liquid coating is applied to a surface with high curvature, such as an airfoil leading edge. The CFD code used, RVCQ3D, is a Reynolds averaged Navier–Stokes solver, with a k-omega turbulence model. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of a coating nonuniformity parameter. Results are presented over a range of negative and positive incidences and inlet Mach numbers from 0.6 to 0.8. A map is provided that shows the allowable degree of coating nonuniformity as a function of incidence and inlet Mach number.

Reduction of Unsteady Blade Loading by Beneficial Use of Vortical and Potential Disturbances in an Axial Compressor With Rotor Clocking

This paper demonstrates reduction of stator unsteady loading due to forced response in a large-scale, low-speed, rotor/stator/rotor axial compressor rig by clocking the downstream rotor. Data from the rotor/stator configuration showed that the stator response due to the upstream vortical disturbance reaches a maximum when the wake impinges against the suction surface immediately downstream of the leading edge. Results from the stator/rotor configuration revealed that the stator response due to the downstream potential disturbance reaches a minimum with a slight time delay after the rotor sweeps pass the stator trailing edge. For the rotor/stator/rotor configuration, with Gap1 = 10 percent chord and Gap2 = 30 percent chord, results showed a 60 percent reduction in the stator force amplitude by clocking the downstream rotor so that the time occurrence of the maximum force due to the upstream vortical disturbance coincides with that of the minimum force due to the downstream potential disturbance. This is the first time, the authors believe, that beneficial use of flow unsteadiness is definitively demonstrated to reduce the blade unsteady loading.

Experimental investigations of the three-dimensional flow in a large deflection turbine cascade with tip clearance

Results obtained from an experimental study of the three-dimensional flow survey within and exit of a large deflection linear turbine cascade are presented for a tip clearance levels of 0.08, 1.5, 3.0 percent of chord and compared with the help of boundary layer probes and that within and exit of a blade passage was done with a miniaturised five hole probe. End wall and blade tip surface static pressures were also obtained, in addition to flow visualisation studies.

Variations in Upstream Vane Loading With Changes in Back Pressure in a Transonic Compressor

Dynamic loading of an inlet guide vane (IGV) in a transonic compressor is characterized by unsteady IGV surface pressures. These pressure data were acquired for two spanwise locations at a 105 percent speed operating condition, which produces supersonic relative Mach numbers over the majority of the rotor blade span. The back pressure of the compressor was varied to determine the effects from such changes. Strong bow shock interaction was evident in both experimental and computational results. Variations in the back pressure have significant influence on the magnitude and phase of the upstream pressure fluctuations. The largest unsteady surface pressure magnitude, 40 kPa, was obtained for the near-stall mass flow condition at 75 percent span and 95 percent chord. Radial variation effects caused by the spanwise variation in relative Mach number were measured. Comparisons to a two-dimensional nonlinear unsteady blade/vane Navier–Stokes analysis show good agreement for the 50 percent span results in terms of IGV unsteady surface pressure. The results of the study indicate that significant nonlinear bow shock influences exist on the IGV trailing edge due to the downstream rotor shock system. [S0889-504X(00)00303-2]

Navier–Stokes and Potential Calculations of Axial Spacing Effect on Vortical and Potential Disturbances and Gust Response in an Axial Compressor

The effect of blade row axial spacing on vortical and potential disturbances and gust response is studied for a compressor stator/rotor configuration near design and at high loadings using two-dimensional incompressible Navier–Stokes and potential codes, both written for multistage calculations. First, vortical and potential disturbances downstream of the isolated stator in the moving frame are defined; these disturbances exclude blade row interaction effects. Then, vortical and potential disturbances for the stator/rotor configuration are calculated for axial gaps of 10, 20, and 30 percent chord. Results show that the potential disturbance is uncoupled locally; the potential disturbance calculated from the isolated stator configuration is a good approximation for that from the stator/rotor configuration upstream of the rotor leading edge at the locations studied. The vortical disturbance depends strongly on blade row interactions. Low-order modes of vortical disturbance are of substantial magnitude and decay much more slowly downstream than do those of potential disturbance. Vortical disturbance decays linearly with increasing mode except very close to the stator trailing edge. For a small axial gap, e.g., 10 percent chord, both vortical and potential disturbances must be included to determine the rotor gust response.

Effect of Angle of Attack on Roll Characteristics of 65-Degree Delta Wing

Results are presented of an analysis that defines the flow physics behind the experimentally observed dramatic change of the roll-trim characteristics of a 65-deg delta wing when the inclination of the roll axis is increased from 30 to 35 deg. The big difference between the Cl(phi) characteristics at sigma = 30 and 35 deg is shown to be a direct consequence of where vortex breakdown occurs at phi = 0; 60 percent chord for sigma = 30 deg compared for 25 percent chord for sigma = 35 percent. Thus, in the former case the vortex-induced lift forward of breakdown dominates, generating a stable trim point at phi = 0, whereas in the latter case the movement of vortex breakdown creates the dominate loads, generating a statically unstable Cl(phi) slope at phi = 0. (Author)

Inverse Design of Centrifugal Compressor Vaned Diffusers in Inlet Shear Flows

A three-dimensional inverse design method in which the blade (or vane) geometry is designed for specified distributions of circulation and blade thickness is applied to the design of centrifugal compressor vaned diffusers. Two generic diffusers are designed, one with uniform inlet flow (equivalent to a conventional design) and the other with a sheared inlet flow. The inlet shear flow effects are modeled in the design method by using the so-called “Secondary Flow Approximation” in which the Bernoulli surfaces are convected by the tangentially mean inviscid flow field. The difference between the vane geometry of the uniform inlet flow and nonuniform inlet flow diffusers is found to be most significant from 50 percent chord to the trailing edge region. The flows through both diffusers are computed by using Denton’s three-dimensional inviscid Euler solver and Dawes’ three-dimensional Navier–Stokes solver under sheared in-flow conditions. The predictions indicate improved pressure recovery and internal flow field for the diffuser designed for shear inlet flow conditions.

Tip Leakage Flow in Linear Compressor Cascade

Tip leakage flow in a linear compressor cascade of NACA 65-1810 profiles is investigated, for tip clearance levels of 1.0, 2.0, and 3.25 percent of chord at design and off-design flow conditions. Velocity and pressure data are collected from three transverse sections inside tip clearance and sixteen sections within flow passage. Tip separation vortex influence is identified from the data. Leakage flow mixing is clearly present inside the clearance and has a significant influence on the internal loss.

Experimental Study on the Three-Dimensional Flow Within a Compressor Cascade With Tip Clearance: Part I—Velocity and Pressure Fields

Experimental results from a study of the three-dimensional flow in a linear compressor cascade with stationary endwall at design conditions are presented for tip clearance levels of 1.0, 2.0, and 3.3 percent of chord, compared with the no-clearance case. In addition to five-hole probe measurements, extensive surface flow visualizations are conducted. It is observed that for the smaller clearance cases a weak horseshoe vortex forms in the front of the blade leading edge. At all the tip gap cases, a multiple tip vortex structure with three discrete vortices around the midchord is found. The tip leakage vortex core is well defined after the midchord but does not cover a significant area in traverse planes. The presence of the tip leakage vortex results in the passage vortex moving close to the endwall and the suction side.

Low Aspect Ratio Transonic Rotors: Part 2—Influence of Location of Maximum Thickness on Transonic Compressor Performance

Transonic compressor rotor performance is sensitive to variations in several known design parameters. One such parameter is the chordwise location of maximum thickness. This article reports on the design and experimental evaluation of two versions of a low aspect ratio transonic rotor that had the location of the tip blade section maximum thickness moved forward in two increments from the nominal 70 percent to 55 and 40 percent chord length, respectively. The original hub characteristics were preserved and the maximum thickness location was adjusted proportionately along the span. Although designed to satisfy identical design speed requirements, the experimental results reveal significant variation in the performance of the rotors. At design speed, the rotor with its maximum thickness located at 55 percent chord length attains the highest peak efficiency among the three rotors but has lowest flow rollback relative to the other two versions. To focus on current ruggedization issues for transonic blading (e.g., bird and ice ingestion), detailed comparison of test data and analysis to characterize the aerodynamic flow details responsible for the measured performance differences were confined to the two rotors with the most forward location of maximum thickness. A three-dimensional viscous flow analysis was used to identify the performance-enhancing features of the higher efficiency rotor and to provide guidance in the interpretation of the experimental measurements. The computational results of the viscous analysis show that the difference in performance between the two rotors can be attributed to the higher shock losses that result from the increased leading edge “wedge angle” as the maximum thickness is moved closer to the leading edge. The test data and the three-dimensional viscous analysis also reveal that the higher efficiency rotor achieves the same static pressure rise potential and loading at a higher flow level than its less efficient counterpart and this is responsible for its resulting lower flow rollback and apparent loss in stall margin. Comparison of the peak efficiencies attained by the two rotors described in this article with the baseline ruggedized rotor performance presented in part 1 of this paper suggests the existence of an optimum maximum thickness location at 55 to 60 percent chord length for such low aspect ratio transonic rotors.

Measurements in a leading-edge separation bubble due to a simulated airfoil ice accretion

The separation bubble formed on an airfoil at low Reynolds number behind a simulated leading-edge glaze ice accretion is studied experimentally. Surface pressure and split hot-film measurements as well as flow visualization studies of the bubble reattachment point are reported. The simulated ice generates an adverse pressure gradient that causes a laminar separation bubble of the long bubble type to form. The boundary layer separates at a location on the ice accretion that is independent of angle of attack and reattaches at a downstream location 5-40 percent chord behind the leading edge, depending on the angle of attack. Velocity profiles show a large region of reverse flow that extends up from the airfoil surface as much as 2.5 percent chord. After reattachment, a thick distorted turbulent boundary layer exists. The separation bubble growth and reattachment are clearly seen in the plots of boundary-layer momentum thickness vs surface distance. Local minima and maxima in the boundary-layer momentum thickness development compare well with the shear layer transition point as indicated by the surface pressures and the reattachment point as measured from surface oil flow, respectively.

A Marked Change in Subcavitating-Cascade-Performance Resulted from a Very Small Modification in Leading Edges.

For the first step to find low-drag thin hydrofoils in cascade applicable to Kaplan-turbines, the subcavitating performance and the corresponding cavitation aspects for two different hydrofoils whose shapes are quite the same except the neighborhood of the leading edge within one percent chord. The leading edges are somewhat modified, for two cascade arrangements of a pitch-chordratio t/l=1.14 and of two cascade angles θ1=23°and 35°, for various cavitation numbers and various incidences. Clearly, a marked change in drag is resulted from such a very small modification around the edge, which is closely related to the boundary layer separation around the leading edge. It is also suggested to the suitable modifications without separation bubble.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 4—Flow Fields for Incidence Angles of −1.5 and −8.5 Degrees

Measurements, made with laser Doppler velocimetry, about a double-circular-arc compressor blade in cascade are presented for −1.5 and −8.5 deg incidence angles and a chord Reynolds number near 500,000. Comparisons between the results of the current study and those of our earlier work at a 5.0 deg incidence are made. It is found that in spite of the relative sophistication of the measurement techniques, transition on the pressure surface at the −1.5 deg incidence is dominated by a separation “bubble” too small to be detected by the laser Doppler velocimeter. The development of the boundary layers at −1.5 and 5.0 deg is found to be similar. In contrast to the flow at these two incidence angles, the leading edge separation bubble is on the pressure surface for the −8.5 deg incidence. Here, all of the measured boundary layers on the pressure surface are turbulent—but extremely thin—while on the suction surface, a laminar separation/turbulent reattachment bubble lies between roughly 35 percent and 60 percent chord. This bubble is quite thin, and some problems in interpreting the backflow data are discussed.

Effects of Shock Waves on Aerodynamic Instability of Annular Cascade Oscillation in a Transonic Flow

The effects of shock waves on the aerodynamic instability of annular cascade oscillation were examined for rows of both turbine and compressor blades, using a controlled-oscillating annular cascade test facility and a method for accurately measuring time-variant pressures on blade surfaces. The nature of the effects and blade surface extent affected by shock waves were clarified over a wide range of Mach number, reduced frequency, and interblade phase angle. Significant unsteady aerodynamic forces were found generated by shock wave movement, which markedly affected the occurrence of compressor cascade flutter as well as turbine cascade flutter. For the turbine cascade, the interblade phase angle significantly controlled the effect of force, while for the compressor cascade the reduced frequency controlled it. The chordwise extent of blade surface affected by shock movement was estimated to be approximately 6 percent chord length.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 2—Suction Surface Boundary Layers

Using the facility described in Part 1 [29], eleven detailed velocity and turbulence intensity profiles are obtained on the suction surface of a double circular arc blade in cascade. At the measured incidence angle of 5 deg, transition through a leading edge separation bubble occurs before 2.6 percent chord. A continuing recovery from this leading edge separation is apparent in the measured boundary layer profiles at 2.6 and 7.6 percent chord. Recovery appears to be complete by 12.7 percent chord. The data then illustrate the evolution of the nonequilibrium turbulent boundary layers as they approach a second region of separation. Following the criteria established by Simpson et al. [1], we find that intermittent separation occurs near 60 percent chord while detachment occurs at 84.2 percent chord. Comparison between the measured profiles and the sublimation visualization studies indicates that the flow visualization is signaling the location of incipient detachment (1 percent instantaneous backflow). Measured profiles are also considered in light of similarity techniques for boundary layers approaching separation. Outer region similarity is shown to vanish for profiles downstream of detachment.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 3—Pressure Surface Boundary Layers and the Near Wake

Using the facility described in Part 1 [23], 11 detailed velocity and turbulence intensity profiles are obtained on the pressure surface of a double circular arc compressor blade in cascade. Two profiles are obtained in the near wake. Laminar boundary layer profiles, which agree well with profiles calculated from Falkner–Skan theory at the local pressure gradient, persist through 57.2 percent chord. The measurements indicate that the onset of transition occurs near 60 percent chord—a value in good agreement with the sublimation flow visualization studies (see Part 1). The lack of a logarithmic region in the data measured at the last chord position (97.9 percent chord) indicates that transition is not complete. The thin laminar boundary layers near the leading edge lead to some measurement problems, which are characterized by large turbulence intensities, in using the laser-Doppler velocimeter (LDV). Close examination of this problem shows that a combination of velocity-gradient broadening and a vibration of the LDV measurement volume causes an elevation of the measured turbulence levels. Fortunately only small errors in mean velocity are introduced. Because of the detached boundary layer on the suction surface, both of the near-wake velocity profiles exhibit regions of backflow. As expected, these near-wake velocity profiles do not exhibit similarity when tested against criteria derived for the far wake.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 1—A Unique Experimental Facility

A unique cascade facility is described which permits the use of laser-Doppler velocimetry (LDV) to measure blade boundary layer profiles. Because of the need for a laser access window, the facility cannot reply on continuous blade pack suction to achieve two-dimensional, periodic flow. Instead, a strong suction upstream of the blade pack is used in combination with tailboards to control the flow field. The distribution of the upstream suction is controlled through a complex baffling system. A periodic, two–dimensional flow field is achieved at a chord Reynolds number of 500,000 and an incidence angle of 5 deg on a highly loaded, double circular arc, compressor blade. Inlet and outlet flow profiles, taken using five-hole probes, and the blade static-pressure distribution are used to document the flow field for use with the LDV measurements (see Parts 2 and 3). Inlet turbulence intensity is measured, using a hot wire, to be 0.18 percent. The static-pressure distribution suggests both separated flow near the trailing edge of the suction surface and an initially laminar boundary layer profile near the leading edge of the pressure surface. Probe measurements are supplemented by sublimation surface visualization studies. The sublimation studies place boundary layer transition at 64.2 ± 3.9 percent chord on the pressure surface, and indicate separation on the suction surface at 65.6 percent ± 3.5 percent chord.