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Laser window design for the army hypersonic compact kinetic energy missile

JOURNAL OF SPACECRAFT AND ROCKETS 2006年第5期43卷 982-989页

作者： Russell, Gerald Cayson, Stephen Jones, Mike Carriger, Wendy Mitchell, Robert Strobel, Forrest Rembert, Michael Gibson, Dave USA Aviat & Missile Res Dev & Engn Ctr AMSRD AMR PS PI Redstone Arsenal AL 35898 USA USA Aviat & Missile Res Dev & Engn Ctr AMSRD AMR SG NC Redstone Arsenal AL 35898 USA ITT Ind Adv Engn Sci Huntsville AL 35806 USA Summit Res Corp Engn Dev Huntsville AL 35802 USA

The U.S. Army is currently developing the compact kinetic energy missile, which achieves hypersonic velocities at sea level. A state-of-the-art guidance technique called the side-scatter laser beam rider has been developed that utilizes a 1.06-mu m laser to deliver an off-axis guidance link to the missile. Four rectangular side-looking optical windows are used on board the missile for the receivers. The peak flight velocity induces severe highly transient thermal gradients. To assess the effect of these gradients, detailed aerothermal and structural analysis followed by ground test experimental validation was conducted. The resulting window design utilized a Dynasil window (fused silica material fabricated by Meller Optics) along with Grafoil to constrain the window and eliminate gas paths to the laser receiver electronics as well as reduce dynamic effects with the metallic holder. The window design was then successfully validated during aerothermal testing conducted at the Naval Air Warfare Center T-Range Aerothermal test Facility and ultimately flight-tested at Eglin Air Force Base. This paper will present a laser window design development process completed through flight testing to demonstrate window system performance.

关键词： Lasers Compact Kinetic Energy Missile flight test programme Air Force Base flight Trajectory Naval Air Warfare Center Structural Analysis Windowing System Thermal Response Compressive Stress

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flight LOADS test OF THE INDIGENOUS DEFENSIVE FIGHTER

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JOURNAL OF AIRCRAFT 1994年第4期31卷 946-952页

作者： CHANG, YB LIN, TY WANG, JR CHEN, CC Aeronautical Research Lab. Taichung Taiwan

flight loads test is a primary structural verification test in fighter aircraft development. The indigenous defensive fighter flight loads test is the first complete flight loads test executed in Taiwan, Republic of China. The basic configuration part of the flight test, including flight loads survey and structural demonstration, was accomplished in 42 flights with external store configurations to follow. Results give both similarity and deviation compared to predictions, and numerous valuable experiences were accumulated from ground calibration test and flight test. One significant improvement in flight test, the automatic roll capability, was introduced to greatly improve maneuver quality during testing and resulted in safe and efficient night test.

关键词： Indigenous Defensive Fighter flight test programme Fighter Aircraft Pitching Moment Digital flight Control Computer Control Surfaces Structural Load Strain Gauge Aircraft Configurations Angle of Attack

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Five-hole flow angle probe calibration from dynamic and tower flyby maneuvers

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JOURNAL OF AIRCRAFT 2005年第1期42卷 80-86页

作者： Parameswaran, V Jategaonkar, RV Press, M Natl Aerosp Labs Flight Mech & Control Div Bangalore 560017 Karnataka India DLR German Aerosp Ctr Inst Flight Res D-38108 Braunschweig Germany

flight-test and data analysis techniques applied to calibrate the static pressure measured by pitot static systems and the flow angles measured by a five-hole probe mounted on a noseboom are described. Dynamic maneuvers with rapid variations in the aircraft motion are analyzed by application of parameter estimation techniques based on the output error method to calibrate the angle of attack and angle of sideslip. A complementary approach based on the Kalman filter technique is applied to a wind-box maneuver to calibrate the flow variables. The tower flyby maneuvers are analyzed using the classical approach of altitude determination. through geometrical evaluation of photographs, and accurate information is derived from redundant sources to calibrate the pitot static system. The investigations showed that the static pressure measured by aircraft-installed pitot static system is accurate, sufficiently whereas that measured by an additional sensor mounted on the noseboom required speed-dependent correction. The flight estimated sensitivity factors for the flow angles measured by the five-hole probe agreed reasonably well with the manufacturer's specifications subject to corrections for biases resulting from misalignment and time delays caused by the recording equipment.

关键词： Angle of Attack Static Pressure Angle of Sideslip Kalman Filter flight test programme Pitot Static System Data Analysis Techniques Sensors Flow Variables Digital Air Data Computer

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DEVELOPMENT AND flight test OF A DEPLOYABLE PRECISION LANDING SYSTEM

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JOURNAL OF AIRCRAFT 1994年第5期31卷 1101-1108页

作者： SIM, AG MURRAY, JE NEUFELD, DC REED, RD NASA Dryden Flight Research Facility Edwards California 93534 PRC Inc. Edwards California 93534

A joint NASA Dryden flight Research Facility and Johnson Space Center program was conducted to determine the feasibility of the autonomous recovery of a spacecraft using a ram-air parafoil system for the final stages of entry from space that included a precision landing. The feasibility of this system was studied using a flight model of a spacecraft in the generic shape of a flattened biconic that weighted approximately 150 lb and was flown under a commerically available, ram-air parachute. Key elements of the vehicle included the Global Positioning System guidance for navigation, flight control computer, ultrasonic sensing for terminal altitude, electronic compass, and onboard data recording. A flight test program was used to develop and refine the vehicle. This vehicle completed an autonomous flight from an altitude of 10,000 ft and a lateral offset of 1.7 miles that resulted in a precision flare and landing into the wind at a predetermined location. At times, the autonomous flight was conducted in the presence of winds approximately equal to vehicle airspeed. Several novel techniques for computing the winds postflight were evaluated. Future program objectives are also presented.

关键词： flight test programme flight Control Computer GPS High Speed flight Parafoil NASA Dryden flight Research Facility Autonomous flight Computing Data Recording Ultralight Airplane

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AIRCRAFT AUTOMATIC flight CONTROL-SYSTEM WITH MODEL INVERSION

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 1987年第3期10卷 269-275页

作者： SMITH, GA MEYER, G NASA AMES RES CTR MOFFETT FIELD CA 94035 USA

Prime carcasses produced loin and round steaks that were more palatable (P < 0.05) than were those from carcasses of Choice through Canner (7 grades) in 85.7% of comparisons and from carcasses of Choice through Standard (3 grades) in 69.0% of comparisons;comparable percentages were 71.4% (6 grades) and 42.9% (2 grades) for Choice and 74.3% (5 grades) and 35.7% (1 grade) for Good. Among Prime through Standard carcasses, grade predicted flavor, tenderness and overall palatability of loin steaks with 30 to 38% accuracy, but could explain no more than 8% of the variation in panel ratings/shear force values of round steaks.

关键词： Automatic flight Control Model Aircraft Elevator Deflection flight test programme Subsonic Speed Helicopters Transient Response Takeoff and Landing Aircraft Configurations Roll Control

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flight-test Experiment Design for Characterizing Stability and Control of Hypersonic Vehicles

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2009年第3期32卷 949-959页

作者： Morelli, Eugene A. NASA Langley Res Ctr Dynam Syst & Control Branch Hampton VA 23681 USA

A maneuver design method that is particularly well suited for determining the stability and control characteristics of hypersonic vehicles is described in detail. Analytical properties of the maneuver design are explained. The importance of these analytical properties for maximizing information content in flight data is discussed, along with practical implementation issues. Results from flight tests of the X-43A hypersonic research vehicle (also called Hyper-X) are used to demonstrate the excellent modeling results obtained using this maneuver design approach. A detailed design procedure for generating the maneuvers is given to allow application to other flight-test programs.

关键词： flight test programme Hypersonic Vehicles Aerodynamic Stability Control Surfaces flight Data Analysis Frequency Domain Elevator Deflection Feedback Control System Hypersonic flight Dynamic Response

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Quiet Tunnel Measurements of HIFiRE-5 Boundary-Layer Transition

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AIAA JOURNAL 2015年第4期53卷 832-846页

作者： Juliano, Thomas J. Borg, Matthew P. Schneider, Steven P. US Air Force Res Lab Wright Patterson AFB OH 45433 USA Purdue Univ W Lafayette IN 47907 USA

The Hypersonic International flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratory and Australian Defence Science and Technology Organisation. The principal goal of HIFiRE flight five is to measure hypersonic boundary-layer transition on a three-dimensional body. The HIFiRE flight tests are supported by a ground test campaign;this paper presents measurements of heat flux and boundary-layer transition in the Boeing/AFOSR Mach-6 Quiet Tunnel. This facility has been developed to provide quiet flow at high Reynolds number, with low noise levels comparable to flight. This tunnel's good optical access enabled global measurement of the heat flux by means of temperature-sensitive paint. Two modes of transition were observed: transition along the centerline, suspected to arise from the amplification of second-mode waves in the boundary layer, and transition roughly halfway between the centerline and leading edges, probably due to the breakdown of crossflow vortices. Reducing the noise level from conventional to quiet substantially delayed both transition modes. Increasing the angle of attack from 0 to 4deg delayed the crossflow transition mode on the windward side. Transition moved forward as freestream unit Reynolds number increased from 2.6.10(6)/m to 11.9.10(6)/m.

关键词： Boundary Layer Separation Heat Flux Distribution Freestream Mach Number Angle of Attack Gortler Vortices flight test programme Leading Edges Air Force Research Laboratory Temperature Sensitive Paint Boeing

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IN-flight SOURCE NOISE OF AN ADVANCED LARGE-SCALE SINGLE-ROTATION PROPELLER

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JOURNAL OF AIRCRAFT 1993年第6期30卷 918-926页

作者： WOODWARD, RP LOEFFLER, IJ NASA Lewis Research Center Cleveland Ohio 44135

A large-scale advanced single-rotation turboprop engine was installed on the left wing of a Gulfstream II aircraft for in-night aeroacoustic tests. This program, designated propfan test assessment (PTA), involved aeroacoustic tests of the propeller over a range of flight conditions. Data was taken both near the source propeller at flight conditions and on the ground, resulting in a unique set of data which is valuable for evaluating acoustic propagation models for cruise noise ground measurements. The in-flight data reported herein was taken for seven test cases. An acoustically instrumented Learjet was flown in formation with the Gulfstream II to acquire noise measurements, and acoustic data was also acquired on the Gulfstream II aircraft. These acoustic measurements defined source levels and directivities for input into long-distance propagation models to predict en route noise. The sideline tone directivities measured by the Learjet showed maximum levels near 105 deg from the propeller upstream axis. Azimuthal directivities based on the maximum observed sideline tone levels showed highest levels below the aircraft (with a + 3-deg propeller axis angle of attack). An investigation of the effect of propeller tip speed (with other engine parameters, such as thrust, shaft power, flight speed, and altitude, held constant) showed that the tone level reduction associated with reductions in propeller tip speed is more significant in the horizontal plane than below the aircraft.

关键词： Single Rotation Propellers Angle of Attack Acoustic Propagation Propfan Turboprop Engine flight Data Acoustic Measurement Aircraft Structures Wing Tip flight test programme

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AERODYNAMICS AND ACOUSTICS OF ROTOR BLADE-VORTEX INTERACTIONS

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JOURNAL OF AIRCRAFT 1995年第5期32卷 970-977页

作者： YU, YH TUNG, C GALLMAN, J SCHULTZ, KJ VANDERWALL, B SPIEGEL, P MICHEA, B NASA AMES RES CTR AEROLIGHTDYNAM DIRECTORATE DIV FLUID MECH MOFFETT FIELD CA 94035 USA

As part of an international cooperative program, several aerodynamic and acoustic prediction codes have been recently developed to understand the rotor blade-vortex interaction (BVI) noise. Furthermore, a joint validation activity of these prediction codes was performed with the blade surface pressure and acoustic data taken in the German-Dutch Wind Tunnel with the 1/7-scale model AH-I Operational Loads Survey blades. Careful attention has been given to the detailed flowfield, especially wake structures during BVIs including the miss-distance between vortex and blade, vortex core size, and tip vortex trajectories. The results of flowfield including wake systems from various prediction codes show some quantitative differences, but the acoustic results using these calculated airloads as inputs compare reasonably well with test data.

关键词： Blade Vortex Interaction Aerodynamic Coefficients Acoustics Vortex Strength German Dutch Wind Tunnels Helicopter Rotor Rotor Wake Pressure Transducers Vortex Lattice Method flight test programme

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Vibration prediction for rotor system with faults using coupled rotor-fuselage model

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JOURNAL OF AIRCRAFT 2004年第2期41卷 348-358页

作者： Yang, M Chopra, I Haas, DJ Univ Maryland Alfred Gessow Rotorcraft Ctr Dept Aerosp Engn College Pk MD 20742 USA USN Ctr Surface Warfare Bethesda MD 20817 USA

A coupled rotor-fuselage vibration analysis is formulated to analyze the effect of rotor system faults on fuselage vibrations and rotor-blade displacements in both hover and forward-flight conditions. Two groups of rotor-system faults, adjustable and nonadjustable component faults, are modeled. Results are presented for an SH-60 helicopter and are compared with available flight-test data. A detailed aeroelastic analysis is carried out where each rotor blade is modeled individually such that rotor dissimilarity can be considered. A fuselage NASTRAN model is incorporated and coupled with the rotor model analysis. The centrifugally tuned, hub-mounted bitilar vibration absorbers are also modeled, and the coupled rotor-fuselage-bifilar nonlinear equations are solved using finite element method in space and time. Predicted 1/rev and 4/rev vibrations at pilot and copilot seats showed acceptable trends for magnitude and phase with flight-test data.

关键词： Articulated Rotor Airframes flight test programme Copilot Finite Element Modeling Vibration Suppression Nastran Mathematical Models Rotor Blades Aeroelastic Analysis

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