Trip Strip Research Articles

In Situ Optical Detection for Ultrasonic Characterization of Materials in a Mach 10 Hypersonic Wind Tunnel

We present results of in situ optical detection of ultrasonic waves generated in stainless steel and graphite samples mounted in a flat plate model during hypersonic flow in the 31-Inch Mach 10 wind tunnel at NASA Langley Research Center. Longitudinal waves are excited in the stainless steel and graphite sample inserts by using a contact piezoelectric transducer and the normal displacement on the surface of the sample, exposed to hypersonic fluid flow, is measured optically using a Sagnac interferometer. Measurements are consistent at different Reynolds numbers both with and without a turbulent trip strip present. Additionally, optical detection of laser-generated surface acoustic waves in a stainless-steel sample during flow is presented. These results are a demonstration of laser-based ultrasonics as an in situ material characterization technique in hypersonic flow, enabling potential applications of in situ monitoring of defect initialization and morphology in hypersonic environments.

Effects of Extended Laminar Flow on Wing Buffet-Onset Characteristics

Global instability analysis is used to investigate the effects of extended regions of laminar flow on both unswept and swept infinite-span wings. The formulation is based on the Reynolds-averaged Navier–Stokes equations and differs from earlier studies on fully turbulent flows in the activation of the trip term in the Spalart–Allmaras eddy-viscosity transport equation. The trip term ensures a rapid transition to turbulence at an arbitrary-specified location on the wing surface, analogous a trip strip in an experiment. A parametric study is conducted for the ONERA OAT15A extruded airfoil, as an example. Results show that an extended region of laminar flow leads to a reduction in the critical angle of attack for the buffet onset, as compared with the fully turbulent conditions. All modes of instability show enhanced growth as a result of the laminar flow. However, increased lift (at fixed angle of attack) associated with an extended region of laminar flow results in a higher lift coefficient at buffet onset as compared with the fully turbulent case. Results show that the laminar flow effects on the buffet onset are linked to the suction-side laminar flow and are largely independent of the boundary-layer state on the pressure side of the wing.

Large Eddy Simulations on Fan Shaped Film Cooling Hole With Upstream Boundary Layer Turbulence Effect Generated by Trip Strip

Abstract Large eddy simulations (LESs) on the well-known 7-7-7 fan shaped cooling hole were carried out. Like using a trip strip to create turbulent boundary layer in practical experiments, trip strips with different configurations were placed upstream of the cooling hole to investigate incoming turbulent boundary layer effect on the film cooling flow behavior. Without the trip, horseshoe vortex generated by laminar boundary layer induced laterally discharging cooling flow in the lateral direction. Meanwhile, the induced cooling flow formed high film cooling effectiveness region around the film cooling hole. When the incoming boundary flow was turbulent, laterally discharged cooling flow was influenced by the turbulent boundary layer to dissipate to the main flow and resultant high effectiveness region disappeared. Depending on the trip configuration, quantitative characteristics of boundary layer such as turbulent intensity, momentum thickness, and shape factor were strongly affected. Some trip configurations resulted in fully developed turbulent boundary layer just before leading edge of the film cooling hole. In such cases, distribution of the film cooling effectiveness showed a reasonable agreement with available experimental data where the quantitative properties of the turbulent boundary layer were similar. However, when the trip was located too close to the film cooling hole, the separated and reattached flow did not develop into the stabilized turbulent boundary layer. Then strong turbulence intensity in the main flow boundary layer stimulated break down of the cooling flow vortex structure and early dissipation to the main flow. It resulted in restricted film cooling flow coverage.

Wind tunnel experiments on a NACA 633‐418 airfoil with different types of leading edge roughness

AbstractThe NACA 633‐418 airfoil has been tested in the new Poul la Cour Wind Tunnel at the Technical University of Denmark, Risø campus, to expand the publicly available data of airfoils with leading edge roughness. The airfoil was constructed with an exchangeable leading edge. The clean airfoil showed good agreement with results from other high‐quality wind tunnels. The airfoil was equipped with three heights of zigzag tape, a trip strip, three sandpaper types, and leading edge cavities with different modifications. In general, increasing zigzag tape heights reduce the maximum lift, whereas the increase in drag is less affected by the zigzag tape height. Almost the same reduction in maximum lift is seen for the different sandpaper types. However, the drag increases with the coarseness of the sandpaper. The highest zigzag tape and coarsest sandpaper resulted in the highest penalty on the aerodynamics. A combination of sandpaper and cavities showed the largest reduction in maximum lift and the highest increase in drag.

The effects of cross section variations on the pressure distribution around a long axisymmetric body

In the present study, the impacts of embedding belts are experimentally studied for a supersonic flow field around a long axisymmetric body. Also, the longitudinal and circumferential pressure coefficients along with the boundary layer profile are investigated for −2 to 6 degrees angles of attack. To this end, two conical–cylindrical belts were installed at the middle and end parts of ogival cylinder model. To ensure the turbulent flow around the model, a trip strip causing artificial disturbances is utilized and the acquired results are compared with the acquired results from the model without a trip strip. To study the effects of the cross section variations on the pressure distribution as well as the boundary layer profiles, three different belts with various leading edge angles were installed at different locations along the cylindrical section of the model. These belts significantly affected both the surface pressure distribution and the boundary layer profile. Passing of flow over the belt leads to sudden variations in pressure coefficient distribution on the belt that is due to sudden cross section variations and consequent development of reverse flow or flow separation regions and production of vortices along the flow path. Also, the presence of belt leads to the development of oblique shock wave on the model, which in turn reduces the Mach number downstream the belt. Studying the circumferential pressure distribution reveals that the presence of belt leads to more asymmetric flow downstream the model, which is intensified as the angle of attack increases.

Experimental study of boundary layer in compressible flow using hot film sensors through statistical and qualitative methods

The boundary layer state, laminar, transitional or turbulent, greatly affects an aircraft’s aerodynamic performance, specifically the skin friction force. In the present research, the characteristics of the boundary layer and the transition region have been studied experimentally through the statistical methods (RMS and Skewness), and also qualitative and frequency analysis using hot film sensors output. The tests have been performed at Mach numbers from 0.25 to 0.85 with the maximum Reynolds number of 11.6 million per meter, and the maximum angle of attack of 9° on a supercritical airfoil. The validation of the methods mentioned, which are usually used for incompressible flow, has been investigated for compressible flow with shock waves with detailed assessment. Results have shown that in the shock- free case, the boundary layer regions are identifiable by statistical analysis such as RMS and Skewness, as in incompressible flow. However, in the presence of shock waves, because of sudden, local and instantaneous changes in the output voltage level of hot films, determination of the dominant pattern in the boundary layer becomes impossible using the RMS of the total data. Therefore, first, the signal must be divided into different regions and then, with identifying the cause of voltage fluctuations in each region and its effect on the behavior of the steady state boundary layer, appropriate intervals for RMS calculation can be determined. The frequency analysis is valid and can be used in all conditions tested. In these tests the trip strip performance in forming forced transition was also evaluated (through using hot films and pressure coefficient distribution) and the characteristics of the boundary layer were studied. Results have illustrated that the trip strip, forced transition, compared with the free transition, causes sudden boundary layer turbulence, limitation of the transitional region, a change in the shock structure and pressure distribution.

Reynolds number sensitivity to aerodynamic forces of twin box bridge girder

This paper presents the experimental results for a streamlined twin box girder of 1545m span suspension bridge in order to investigate the Reynolds number sensitivity to aerodynamic force and pressure coefficients. High speed wind tunnel testing was carried out on a 1:30 scale sectional model at an aeronautical wind tunnel. The drag and lift coefficients revealed significant decreases at a critical Reynolds number that was in the range of ordinary wind tunnel tests. The safety rail reduced the Reynolds number dependency of aerodynamic force coefficients at negative angles of attack. Similarly, the boundary layer trip strip attached at the bottom surface of the girder reduced the Reynolds number dependency by fixing the flow separation. The pressure coefficients near the gap between the twin boxes changed from negative to positive at the critical Reynolds number, which represented smooth ventilation of wind flow through the gap at the supercritical region. The imbalance of the drag and lift forces acting on each box in the subcritical region gradually resolved and both boxes were almost equally loaded by the forces at the supercritical region.

Lock-in of elastically mounted airfoils at a 90° angle of attack

Reducing vortex-induced vibration (VIV) of elastically mounted cylinders has applications to petroleum, nuclear, and civil engineering. One simple method is streamlining the cylinder into an airfoil shape. However, if flow direction changes, an elastic airfoil could experience similar oscillations with even more drag. To better understand a general airfoil's response, three elastically mounted airfoil shapes are tested at a 90° angle of attack in a 3ft by 4ft wind tunnel. The shapes are a NACA 0018, a sharp leading- and trailing-edge (sharp–sharp) model, and a round leading- and trailing-edge (round–round) model. Mass-damping ranges from 0.96 to 1.44. For comparison to canonical VIV research, a cylinder is also tested. Since lock-in occurs near Rec=125×103, the models are also tested with a trip strip. The NACA 0018 and sharp–sharp configuration show nearly identical responses. The cylinder and round–round airfoil have responses five to eight times larger. Thus, the existence of a single sharp edge is sufficient to greatly reduce VIV at 90° angle of attack. Whereas the cylinder and round–round maximum response amplitudes are similar, cylinder lock-in occurs over a velocity range three times larger than the round–round. The tripped cylinder and round–round models' response is attenuated by 70% compared to their respective clean configurations. Hysteresis is only observed in the circular cylinder and round–round models. Hotwire data indicates the clean cylinder has a unique vortex pattern compared to the other configurations.

레이놀즈수에 따른 이순신대교 거더에 작용하는 공기력의 변화

본 연구에서는 트윈박스 단면인 이순신대교의 레이놀즈수 변화에 따른 공기력의 영향을 살펴보는데 그 목적이 있다. 이를 위하여 1/30 대축척 모형을 제작하여 공군사관학교 아음속 중형 풍동에서 최대 풍속 70m/s까지 풍속을 증가시켜가면서 공기력을 측정하여, 전북대학교 소형풍동에서 수행한 저레이놀즈수 풍동실험 결과와 비교하였다. 본 연구 대상 교량 단면은 레이놀즈수의 영향을 받는 것으로 나타났으며, 고레이놀즈수 실험 결과 기존 저레이놀즈수 실험보다 항력계수는 약 23%정도 낮은 수준인걸로 나타났다. 또한 경계층 촉진장치를 효과적으로 이용하면 기존의 저레이놀즈수 풍동실험 조건에서 고레이놀즈수 모사 실험이 가능한 것으로 판단된다. The objective of present study is to investigate the sensitivity of aerostatic force coefficients of twin box girder of Yi Sun-sin Bridge according to the Reynolds numbers. This paper presents the 1:30 scale sectional model tests conducted at high speed wind tunnel in Korea Air Force Academy. Comparison with results at low Reynolds number obtained in KOCED Wind Tunnel Center in Chonbuk National University is also provide. The Reynolds number dependency of aerodynamic force coefficients were observed at present streamlined twin box girder. The drag coefficient revealed significant decrease of nearby 23% at supercritical region. The boundary layer trip strip was found to reduce the Reynolds number dependency of aerodynamic forces by fixing the location of flow transition.

Experimental study on correlation between turbulence and sound in a subsonic wind tunnel

In this paper, the effects of turbulence on sound generation and velocity fluctuations due to pressure waves in a large subsonic wind tunnel are studied. A trip strip located at different positions in the contraction part or at one position in the diffuser of a large wind tunnel is used to investigate the aforementioned phenomenon, and the results indicate that the trip strip has significant effects on sound reduction. The lowest turbulence intensity and sound are obtained from a trip strip with a diameter of 0.91 mm located either at X/L = 0.79 or at X/L = 0.115 in the wide portion of the contraction. Furthermore, the effect of monopole, dipole and quadrupole sources of aerodynamic noise at different velocities is investigated, and it is demonstrated that the contribution of the monopole is dominant, while the shares due to the dipole and quadrupole remain less important. In addition, it is found that the sound waves have a modest impact on the measured longitudinal turbulence and are generated essentially by eddies.

Control of separation in the concave portion of contraction to improve the flow quality

AbstractSubsonic wind tunnel experiments were conducted to study the effect of forced transition on the pressure distribution in the concave portion of contraction. Further more, the effect of early transition on the turbulence level in the test section of the wind tunnel is studied. Measurements were performed by installing several trip strips at two different positions in the concave portion of the contraction. The results show that installation of the trip strips, have significant effects on both turbulence intensity and on the pressure distribution. The reduction in the free stream turbulence as well as the wall static pressure distribution varied significantly with the location of the trip strip. The results confirm the significant impact of the tripped boundary layer on the control of adverse pressure gradient. The trip strip at X/L = 0.115 improves pressure distribution in contraction and reduces turbulence intensity in the test section, considerably.

An implicit and explicit BEM sensitivity approach for thermo-structural optimization

A computer-automated shape optimization methodology has been developed for the purpose of providing internal cooling systems designers the ability to optimize the internal cooling configuration, geometry and heat transfer enhancements for greater cooling efficiency and more durable turbine airfoils. The methodology presents the theory and practical programming requirements for coupling existing computer design and analysis tools together into a new and powerful design system. The goal of this paper is to demonstrate the computational advantages of using implicit sensitivity with the boundary element method (BEM) within this system over other more brute force methods. For this research, BEM algorithms for nonlinear heat conduction and thermo-elasticity were developed and coupled to an unstructured finite volume CFD code for the hot gas flow and a quasi-one-dimensional thermo-fluid system for the analysis of the internal coolant network. These computational tools were controlled by a constrained hybrid optimization algorithm to provide aerodynamic, thermal and internal fluid flow analyses on modified designs. The coolant supply total pressure, turbine inlet temperature, coolant wall thickness, thickness of ribs, rib positions, rib orientations, pin fin diameters and trip strip heights were incorporated into the set of optimization design variables. In order to improve performance, sensitivity gradients of the objective and constraint functions with respect to the geometric and heat transfer enhancement design variables were obtained using implicit differentiation of the boundary element system of equations. A three-to-one improvement in the optimization convergence rate and greater gradient accuracy were obtained for the two-dimensional thermal optimization problems. An order of magnitude larger computing time reduction was realized for three-dimensional thermal optimizations at the expense of additional memory, and another order of magnitude is expected for thermo-elastic optimization problems. Examples include studies of the accuracy of the design sensitivities with respect to forward and central finite differences, and validation of the optimization process using a symmetric cooled configuration.

Correlation of Helicopter Rotor Tip Vortex Measurements

The roll-up and development of the tip vortices generated by subscale helicopter rotors were measured and analyzed. The measurements were performed on one- and two-bladed hovering rotors using three-component laser Doppler velocimetry. A trip strip was used to simulate boundary-layer transition at the higher Reynolds numbers typical of the full-scale rotor. The detailed velocity e eld of the tip vortex was investigated, a strong selfsimilarity of the swirl velocity proe le being found. The measurements suggest a general logarithmic growth of the viscous core, which for older vortex ages is proportional to the square root of time (vortex age). Measurements from other helicopter rotor and e xed-wing experiments are used to support the observations.

56 衝撃波管を用いた翼型流れにおける流れの可視化と問題点

56 衝撃波管を用いた翼型流れにおける流れの可視化と問題点

7 - Risk Associations in the Thrombotic Disorders

Reducing vortex-induced vibration (VIV) of elastically mounted cylinders has applications to petroleum, nuclear, and civil engineering. One simple method is streamlining the cylinder into an airfoil shape. However, if flow direction changes, an elastic airfoil could experience similar oscillations with even more drag. To better understand a general airfoil's response, three elastically mounted airfoil shapes are tested at a 90° angle of attack in a 3 ft by 4 ft wind tunnel. The shapes are a NACA 0018, a sharp leading- and trailing-edge (sharp–sharp) model, and a round leading- and trailing-edge (round–round) model. Mass-damping ranges from 0.96 to 1.44. For comparison to canonical VIV research, a cylinder is also tested. Since lock-in occurs near Rec=125×103, the models are also tested with a trip strip. The NACA 0018 and sharp–sharp configuration show nearly identical responses. The cylinder and round–round airfoil have responses five to eight times larger. Thus, the existence of a single sharp edge is sufficient to greatly reduce VIV at 90° angle of attack. Whereas the cylinder and round–round maximum response amplitudes are similar, cylinder lock-in occurs over a velocity range three times larger than the round–round. The tripped cylinder and round–round models' response is attenuated by 70% compared to their respective clean configurations. Hysteresis is only observed in the circular cylinder and round–round models. Hotwire data indicates the clean cylinder has a unique vortex pattern compared to the other configurations.