检索结果-内蒙古大学图书馆

AIAA Science and Technology Forum and Exposition (AIAA SciTech)

作者： Owens, Lewis R. Beeler, George B. King, Rudolph A. Chou, Amanda Balakumar, Ponnampalam Banks, Daniel W. NASA Langley Res Ctr Flow Phys & Control Branch Hampton VA 23681 USA NASA Langley Res Ctr Hampton VA 23681 USA NASA Armstrong Flight Res Ctr Aerodynam & Prop Branch Edwards AFB CA 93523 USA Jacobs Technol Inc Tullahoma TN USA

ISBN: (数字)9781624105951

ISBN: (纸本)9781624105951

This paper continues analysis of data acquired in a series of supersonic crossflow transition experiments on a 65 degrees swept-wing model tested in both wind tunnel and flight. flow visualization data are summarized to provide context for the swept-wing instability mechanisms of interest. Unsteady pressure measurements obtained near the wing model leading edge (LE) are studied using 2nd- and 3rd-order statistical methods to extract traveling crossflow wave characteristics from these data. Comparisons between tunnel and flight data with boundary-layer (BL) Linear Stability Theory (LST) predictions provide insights into BL transition phenomena similarities and differences observed in tests using the same swept-wing model configurations. Comparisons of unsteady pressure data results for wing LE configurations with and without distributed-roughness-element (DRE) patterns installed, suggest traveling crossflow disturbance growth is reduced when this type of flow control is applied. A brief study of different background surface roughness levels on the wing LE suggests the potential for another flow control approach to promote laminar flow on swept wings using streamwise-biased surface finishes.

关键词： Flight Testing Leading Edges Potential flow flow visualization techniques Surface Roughness Boundary Layers Flight Data Supersonic Wind Tunnels Piezoelectric Transducers Computational Fluid Dynamics

来源：评论

学校读者我要写书评

暂无评论

Experimental Aerodynamic Analysis of a 4.6%-Scale Flying-V Subsonic Transport

Experimental Aerodynamic Analysis of a 4.6%-Scale Flying-V S...

引用

AIAA Science and Technology Forum and Exposition (AIAA SciTech)

作者： Palermo, Marco Vos, Roelof Delft Univ Technol Fac Aerosp Engn NL-2629 HS Delft Netherlands

ISBN: (数字)9781624105951

ISBN: (纸本)9781624105951

The experimental investigation into the aerodynamic characteristics of a 4.6%-scale half-model of a flying-wing transport aircraft is detailed. The study is performed in an open-jet wind tunnel where forces and moments are recorded using a 6-axis balance. Two control surfaces in the outboard wing are deflected to measure their effect in terms of lift, drag, and pitching moment up to a Reynolds number of 1 million. The results are subsequently combined with the estimated thrust force using a flight mechanics model of the aircraft in order to predict the most forward and most aft center-of-gravity locations for which the aircraft can be balanced with the control surfaces, while still being statically stable. The results show that the aircraft can attain an untrimmed maximum lift coefficient of 1.02 at an angle of attack of 35 degrees. Furthermore, the pitching moment around the leading-edge of the mean geometric chord is negatively correlated to the angle of attack up to 19 degrees, after which a strong pitch-break is observed, making the aircraft statically unstable. This is associated with a forward shift of the aerodynamic center to a longitudinal position 35%(c) over bar ahead of the moment reference point which is caused by the formation of strong vortices over the wing surface. The effectiveness of the control surfaces hardly deteriorates with angle of attack and all three control surfaces are shown to be effective up to the maximum lift coefficient. Analysis shows that the center-of-gravity location should reside between [-7.5;0.5]%(c) over bar, in power-off conditions, and between [-6;1] %(c) over bar, in power-on conditions, from the leading edge of the mean geometric chord to ensure an ultimate static stability margin of 4.4% as well and a minimum landing speed of 20 m/s. Within these ranges, trimmed maximum lift coefficient values of 0.68 and 0.66 can be achieved respectively in power-off and power-on conditions.

关键词： Aerodynamic Efficiency Aerodynamic Characteristics Control Surfaces Low Speed Wind Tunnel Lift to Drag Ratio Vortex Shedding Transport Aircraft Flight Mechanics Low Aspect Ratio Wings flow visualization techniques

来源：评论

学校读者我要写书评

暂无评论

Design and Analysis of an Additive Manufactured Supersonic Wind Tunnel

Design and Analysis of an Additive Manufactured Supersonic W...

引用

AIAA Science and Technology Forum and Exposition (AIAA SciTech)

作者： Gazella, Matthew R. Hill, Jonathon R. Chriss, Scott L. Abel, Austin J. Gunasekaran, Sidaard US Air Force Res Lab Wright Patterson AFB OH 45433 USA Univ Dayton Dept Mech & Aero Engn Dayton OH 45469 USA

ISBN: (数字)9781624105951

ISBN: (纸本)9781624105951

For decades, supersonic wind tunnels have been designed and fabricated using common metallic materials for structural integrity. With advancements in additive manufacturing technology, high-performance thermoplastics are being developed as an alternative to steel that offer excellent strength and thermal stability. Utilizing the latest advancements in additive manufacturing, an additive manufactured supersonic wind tunnel with Mach 2+ capability was designed and fabricated using ULTEM (TM) 1010 resin for educational and research purposes at the University of Dayton Aerospace Laboratory. The supersonic wind tunnel features an open loop, blowdown design utilizing a high-pressure nitrogen cylinder with a settling chamber, modular C-D facility nozzle, modular 1" x 1" test section with optical access for flow visualization, and a step diffuser. High-fidelity CFD analysis was performed using CFD++ software on an AFRL supercomputer which verified the flowpath design estimated from method of characteristics and ideal operating conditions. The CFD analysis was performed with over 10.5 million cells and assumed compressible perfect gas while using a RANS 2-equation cubic k-epsilon turbulence model with viscous terms. Experimental tests were conducted using a Schlieren flow visualization system to validate the CFD analysis and characterize operating conditions for future educational and research activities.

关键词： Supersonic Wind Tunnels flow visualization techniques CFD Analysis Thermoplastic Air Force Research Laboratory Open Loop Thermal Stability Supercomputers Structural Integrity Perfect Gas

来源：评论

学校读者我要写书评

暂无评论

Prediction of Dual-Rotor-Wing Interaction in Hover Using Actuator Disk Theory

Prediction of Dual-Rotor-Wing Interaction in Hover Using Act...

引用

AIAA Aviation Forum

作者： Chen, Mingtai Hubner, James P. Univ Alabama Dept Aerosp Engn & Mech Tuscaloosa AL 35487 USA

ISBN: (数字)9781624105982

ISBN: (纸本)9781624105982

This paper details a potential flow solution for dual-rotor-wing interactions based on actuator disk theory, the method of images and momentum theory in terms of rotor thrust and wing download force of small tiltrotor-like vehicles. The regime of interest is Reynolds number < 100,000 and disk loading < 100 N/m(2). The analytical model examines the effect of rotor height above the wing, the rotor disk coverage over the wing, and the spacing between two rotors. The analytical model requires the rotor radius, which assists in preliminary design studies on tiltrotor-like vehicles. The analytical results compare well with the experimental data in terms of magnitude and trend, particularly the rotor thrust.

关键词： Blade Loading flow visualization techniques Counter Rotating Rotor Biot Savart Law Incompressible flow Vortex Sheet Thrust Augmentation Mass flow Rate Air Vehicle MATLAB

来源：评论

学校读者我要写书评

暂无评论

Experimental study of a CFD validation test case for turbulent separated flows

Experimental study of a CFD validation test case for turbule...

引用

AIAA Science and Technology Forum and Exposition (AIAA SciTech)

作者： Williams, Owen Samuell, Madeline Sarwas, Sage Robbins, Matthew Ferrante, Antonino Univ Washington William E Boeing Dept Aeronaut & Astronaut Seattle WA 98195 USA

ISBN: (数字)9781624105951

ISBN: (纸本)9781624105951

An experimental and computational investigation of a new turbulent separated flow simulation validation test case at subsonic flow conditions is presented. Through close collaboration with Boeing, a three-dimensional tapered hump or "speed bump" geometry, relevant to a range of aerodynamic applications and central to an integrated testing/CFD campaign to improve turbulent RANS modeling is proposed. This study focuses on an examination of the flow over this geometry, as well as surveys of inflow boundary conditions, allowing the formalization of the domain, geometry and inflow length for corresponding CFD simulations. The analytically-defined geometric shape is tapered at the sides to minimize adverse side-wall flow interactions present in several previous bump experiments. Surface pressure and flow visualizations indicate the presence of moderate three-dimensionality at the centerline of the model but also a regime of Reynolds number independence, which is a desirable property for a validation test case. In contrast, while pressure coefficients computed from RANS simulations have similar magnitude to the experimental results, they fail to capture the Reynolds number insensitivity and profile inflection near reattachment that is observed in the experiments. These results suggest that the chosen geometry provides a strongly separated flow that is a challenging test for current RANS simulation capabilities.

关键词： Transitional flow Freestream Conditions flow visualization techniques CFD Simulation Boeing Boundary Layer Thickness flow Characteristics Wind Tunnel Tests Vortices Spalart Allmaras Turbulence Model

来源：评论

学校读者我要写书评

暂无评论

Evidence of Surface Curvature Effects in Smooth Body flow Separation Experiments

Evidence of Surface Curvature Effects in Smooth Body Flow Se...

引用

AIAA Aviation Forum and Exposition

作者： Simmons, Daniel J. Thomas, Flint O. Corke, Thomas C. Inst Flow Phys & Control Dept Aerosp & Mech Engn Notre Dame IN 46556 USA

ISBN: (数字)9781624105890

ISBN: (纸本)9781624105890

This paper describes experiments on smooth body adverse pressure gradient (APG) turbulent boundary layer flow separation for the purpose of CFD validation and model development. The experimental geometry is designed to provide initial canonical turbulent boundary layer growth under nominally zero pressure gradient conditions prior to encountering a large, smooth, two-dimensional convex ramp geometry onto which a streamwise APG that is fully adjustable is imposed. Experiments encompassing both largeand small-scale smooth body flow separation as well as attached APG turbulent boundary layer development are realized on the same ramp geometry. The reported experiments are performed at M-infinity = 0.2 which corresponds to Re-theta approximate to 11,000 at the onset of the APG. Sample results are presented for the small-scale separation case and these include the evolution of both the mean flow and turbulent stresses at multiple spanwise locations on the ramp. Application of embedded shear layer scaling is shown to fully collapse the inflectional mean velocity profiles at all streamwise and spanwise locations on the ramp. When similar scaling is applied to the turbulent stresses, evidence of the effect of ramp surface curvature emerges. Due to the rapid transition from initially convex to concave curvature, it is demonstrated that the Reynolds stresses are effectively frozen.

关键词： Boundary Layer Separation flow visualization techniques Velocity Profiles Normal Stress Wind Tunnel Diffuser flow Conditions Freestream Velocity Laser Doppler Velocimetry Freestream Turbulence Static Pressure

来源：评论

学校读者我要写书评

暂无评论

Experimental Investigation of the NASA Common Research Model with a Natural Laminar flow Wing in the NASA Langley National Transonic Facility

Experimental Investigation of the NASA Common Research Model...

引用

AIAA Science and Technology (SciTech) Forum and Exposition

作者： Rivers, Melissa B. Lynde, Michelle Campbell, Richard Viken, Sally Chan, David Watkins, A. Neal Goodliff, Scott NASA Langley Res Ctr Configurat Aerodynam Branch Mail Stop 267 Hampton VA 23681 USA NASA Langley Res Ctr Configurat Aerodynam Branch Mail Stop 499 Hampton VA 23681 USA NASA Langley Res Ctr Mail Stop 493 Hampton VA 23681 USA Jacobs Technol Inc Hampton VA 23681 USA

ISBN: (数字)9781624105784

ISBN: (纸本)9781624105784

A test of the new NASA Common Research Model with a Natural Laminar flow (CRM-NLF) semispan wing in the NASA Langley National Transonic Facility (NTF) was completed in October 2018. The main focus of this test was the evaluation of the extent of laminar flow on the CRM-NLF wing at various Reynolds numbers and test conditions. During this test, data were acquired at chord Reynolds numbers from 10 to 30 million and at Mach numbers ranging from 0.84 to 0.86. This investigation provided valuable insight into the necessary procedures for laminar flow testing in the NTF. It also significantly advanced the new carbon-based heating layer technique to improve the quality of transition visualization data from temperature sensitive paint (TSP) in a cryogenic wind tunnel.

关键词： NASA Langley Natural Laminar flow National Transonic Facility Freestream Mach Number Cryogenic Wind Tunnel Temperature Sensitive Paint Aircraft Configurations Angle of Attack Boundary Layer Transition flow visualization techniques

来源：评论

学校读者我要写书评

暂无评论

Study on the Effectiveness of Plasma Induced flow in Manipulating a Near-Wall Vortex

Study on the Effectiveness of Plasma Induced Flow in Manipul...

引用

AIAA Science and Technology (SciTech) Forum and Exposition

作者： Wittig, T. Nino, G. F. Breidenthal, R. E. Tech Univ Berlin D-10623 Berlin Germany Univ Washington Dept Aeronaut & Astronaut Seattle WA 98195 USA

ISBN: (数字)9781624105784

ISBN: (纸本)9781624105784

This paper presents an experimental study on the manipulation of vortical flow with a nanosecond dielectric barrier discharge (ns-DBD) plasma actuator. A vortex generator (VG) with a swept back leading edge was mounted to a flat plate and angled to the free stream to generate a vortex. The angle of attack of the VG ranged from 10-25 degrees with 5 degrees increments. The experiments were performed in the Kirsten Wind Tunnel at the University of Washington with a free stream velocity ranging from 20-70 m/ s. In order to analyze the manipulation of the vortical flow through plasma actuation, an array with 120 pressure ports was installed downstream of the VG. The plasma actuator was mounted near the swept back leading edge of the VG. The experiments reveal that the plasma induced electric wind results in a decrease of the vortex drift angle. It is believed that the induced body force displaces the center line of the induced vortex further away from the wall. Counter-intuitively, this corresponds to a reduction of the lift. Following the Kutta-Joukowski theorem, a reduction of the lift is associated with a reduction of the circulation and the vortex drift angle.

关键词： Karman Vortex Street Lift Coefficient flow visualization techniques Kutta Joukowski Theorem Plasma Actuator Dielectric Barrier Discharge Wind Tunnel Models Pressure Sensors Turbulent Boundary Layer Skin Friction

来源：评论

学校读者我要写书评

暂无评论

Effects of Double-Wedge Aft Angle on Hypersonic Laminar flows

引用

AIAA JOURNAL 2020年第4期58卷 1689-1703页

作者： Durna, A. S. Celik, B. Samsun Univ Dept Aerosp Engn TR-55420 Samsun Turkey Istanbul Tech Univ Dept Astronaut Engn TR-34469 Istanbul Turkey

The aim of this study is to reveal the physics of low-enthalpy Mach 7 flows over double wedges that have varied aft angles and a fixed fore angle of 30 degrees. Comparisons of the obtained numerical schlieren images along with the contours of Q criterion and surface heat fluxes assert that the influence of the vortical structures on the flow becomes more significant in accordance with the value of the aft angle. For the aft angle of theta 2<50 degrees, the three-dimensionality effects are negligible. As the aft angle exceeds 50 degrees, the vortical structures that align in the streamwise direction start to emerge and strengthen over the surfaces of the aft wedge. The characteristic signs of Gortler vortices, distinct striations, also become apparent in the contours of heat flux, density gradient, and pressure on the surfaces of the aft wedge for those cases. The presence of vortical structures enhances heat transfer from the surfaces and together with relieving effects breaks the symmetry down in the flowfield. In general, the vortical structures and relieving effects that take place in the computations diversify the shock structure from those obtained from the two-dimensional computations for the cases with theta 2>50 degrees.

关键词： Compressible flow Heat Flux Distribution Gortler Vortices Surface Heat Transfer Enthalpy flow visualization techniques Boundary Layer Separation Surface Streamlines Navier Stokes Solver Freestream Conditions

来源：评论

学校读者我要写书评

暂无评论

Underexpanded Supersonic Jets from Elliptical Nozzle with Aft Deck

引用

JOURNAL OF PROPULSION AND POWER 2020年第1期36卷 138-152页

作者： Rao, A. Nageswara Kushari, Abhijit Indian Inst Technol Dept Aerosp Engn Kanpur 208016 Uttar Pradesh India

An experimental investigation of an underexpanded supersonic jet issuing from a circular nozzle, an elliptical nozzle, and an elliptical nozzle with an aft deck at different nozzle pressure ratios (NPRs) is reported in this paper. Static-pressure measurements over the nozzle walls were analyzed to understand the internal flow structure. Schlieren visualization and conical probe measurements were conducted to elucidate the development of the jets issuing from the nozzles. The sound-pressure level in the near field was measured, and it was found that the screech amplitudes were 26 dB lower in the elliptical nozzle with an aft deck. The screech-tone frequencies were observed to be above and near audible region (16 kHz) in the presence of the aft deck at NPRs 3 and 4. The jets emanating from the circular and truncated elliptical nozzles were observed to be flapping about the jet axis at NPRs 3 and 4. The unsteadiness of the jet was reduced by locking the jet onto the aft deck. The centerline Mach-number decay revealed that, due to the screeching effect, the truncated elliptical nozzle has a shorter potential core length and larger jet spreading. The flow separation and reattachment were observed over the aft-deck surface for higher NPR conditions.

关键词： Convergent Nozzles Mach Numbers Static Pressure Boundary Layer Separation Sound Pressure Level Compressible flow High Aspect Ratio flow visualization techniques Feedback Loop Oblique Shock

来源：评论

学校读者我要写书评

暂无评论

建议与咨询留下您的常用邮箱和电话号码，以便我们向您反馈解决方案和替代方法

时间限定

文献类型

馆藏选择

核心期刊

语言

文献类型

帮助

文字说明：

检索规则说明：

检索范例：

分类表

所选分类

限定检索结果

文献类型

馆藏范围

日期分布

学科分类号

主题

机构

作者

语言

请选择保存的检索档案：

请选择收藏分类：

通借通还

建议与咨询 留下您的常用邮箱和电话号码，以便我们向您反馈解决方案和替代方法

时间限定

文献类型

馆藏选择

核心期刊

语言

文献类型

帮助

文字说明：

检索规则说明：

检索范例：

分类表

所选分类

限定检索结果

文献类型

馆藏范围

日期分布

学科分类号

主题

机构

作者

语言

请选择保存的检索档案： 新增检索档案 确定 取消

请选择收藏分类： 新增自定义分类 确定 取消

通借通还

建议与咨询留下您的常用邮箱和电话号码，以便我们向您反馈解决方案和替代方法

请选择保存的检索档案：

请选择收藏分类：