Saint Louis, MO (Contract NAS1-18037) (NASA-CR-181817; NAS 1.26:181817) Avail: NTIS HC A06/MF A01 CSCL 01C An evaluation was made of the applicability and benefits of laminar flow control (LFC) technology to supersonic cruise airplanes. Ancillary objectives were to identify the technical issues critical to supersonic LFC application, and to determine how those issues can be addressed through flight and wind-tunnel testing. Vehicle types studied include a Mach 2.2 supersonic transport configuration, a Mach 4.0 transport, and two Mach 2-class fighter concepts. Laminar flow control methodologies developed for subsonic and transonic wing laminarization were extended and applied. No intractible aerodynamic problems were found in applying LFC to airplanes of the Mach 2 class, even ones of large size. Improvements of 12 to 17 percent in lift-drag ratios were found. Several key technical issues, such as contamination avoidance and excresence criteria were identified. Recommendations are made for their resolution. A need for an inverse supersonic wing design methodology is indicated. Author N89-26842*# Douglas Aircraft Co., Inc., Long Beach, CA. COMPOSITE TRANSPORT WING TECHNOLOGY DEVELOPMENT Ram C. Madan Feb. 1988 92 p (Contract NAS1-17970) (NASA-CR-178409; NAS 1.26:178409) Avail: NTIS HC A05/MF A01 CSCL 01C The design, fabrication, testing, and analysis of stiffened wing cover panels to assess damage tolerance criteria are discussed. The damage tolerance improvements were demonstrated in a test program using full-sized cover panel subcomponents. The panels utilized a hard skin concept with identical laminates of 44-percent 0-degree, 44-percent plus or minus 45-degree, and 12-percent 90-degree plies in the skins and stiffeners. The panel skins were impacted at midbay between the stiffeners, directly over the stiffener, and over the stiffener flange edge. The stiffener blades were impacted laterally. Impact energy levels of 100 ft-lb and 200 ft-lb were used. NASTRAN finite-element analyses were performed to simulate the nonvisible damage that was detected in the panels by nondestructive inspection. A closed-form solution for generalized loading was developed to evaluate the peel stresses in the bonded structure. Two-dimensional delamination growth analysis was developed using the principle of minimum potential energy in terms of closed-form solution for critical strain. An analysis was conducted to determine the residual compressive stress in the panels after impact damage, and the analytical predictions were verified by compression testing of the damaged panels. Author N89-26843*# Army Aerostructures Directorate, Hampton, VA. National Technical Specialists' Meeting on Advanced Rotorcraft Structures, Williamsburg, VA, 25-27 Oct. 1988 Previously announced in IAA as A89-29466 (NASA-TM-101619; NAS 1.15:101619; AVSCOM-TM-89-B-006) Avail: NTIS HC A03/MF A01 CSCL 01C of Scaling effects in the large deflection response graphite-epoxy composite beams was investigated. Eight different scale model beams ranging from 1/6 to full-scale were subjected to an eccentric axial compressive load to promote large bending deformations and failures. Beams having laminate stacking sequences including unidirectional, angle ply, cross ply, and quasi-isotropic were tested to examine a wide variety of composite response and failure modes. The model beams were loaded under scaled test conditions until catastrophic failure. Data acquired included load, end displacement, and strain measurements, and qualitative failure measurements. The experimental data is compared to a large rotation beam analysis and a finite element model analysis. Results from the tests indicate that the beam response becomes nonlinear. Failure modes are consistent (NASA-TP-2923; H-1510; NAS 1.60:2923) Avail: NTIS HC A03/MF A01 CSCL 01C A method for flight flutter testing is proposed which enables one to determine the flutter dynamic pressure from flights flown far below the flutter dynamic pressure. The method is based on the identification of the coefficients of the equations of motion at low dynamic pressures, followed by the solution of these equations to compute the flutter dynamic pressure. The initial results of simulated data reported in the present work indicate that the method can accurately predict the flutter dynamic pressure, as described. If no insurmountable difficulties arise in the implementation of this method, it may significantly improve the procedures for flight flutter testing. Author 06 AIRCRAFT INSTRUMENTATION Includes cockpit and cabin display devices; and flight instruments. For related information see also 19 Spacecraft Instrumentation and 35 Instrumentation and Photography. N89-26845 Department of the Navy, Washington, DC. COST-OPTIMAL STATE FEEDBACK CONTROLLER FOR ALL-ATTITUDE GIMBAL SYSTEM Patent Filed George L. Lauro, inventor (to Navy) 7 Mar. 1989 11 P 2 Apr. 1987 (AD-D013972; US-Patent-4,811,233; US-Patent-Appl-SN-036225) Avail: US Patent and Trademark Office CSCL 17/7 An improved cost-optimal state feedback control system is disclosed for stabilizing the inertial reference platform in all-attitude inertial guidance systems which uses the region control concept. It reduces the computational requirements of cost-minimizing state feedback to within the capacity of aerospace vehicle on board control processors. The control system iterates the feedback compensation process so as to reduce the R-gyro error to essentially a zero value and to prevent the gimbal assembly from entering a gimbal lock orientation. GRA F-15E EADI (ELECTRONIC ATTITUDE DIRECTOR INDICATOR) EVALUATION: A COMPARISON OF THREE FORMATS Final Report, Apr.- May 1988 R. K. Burns and P. B. Lovering (Midwest Research Inst., Dayton, OH.) Jun. 1988 74 P (AD-A206809; ASD-TR-88-5030) Avail: NTIS HC A04/MF A01 CSCL 01/3 The primary task for any pilot is to maintain attitude awareness and spatial orientation throughout all phases of flight. Even the slightest hesitation or attitude interpretation error during the all intensive low altitude, high speed night mission could rapidly become a catastrophic error. The attitude display must provide an unambiguous presentation of aircraft attitude in a manner which allows the pilot to respond immediately without thinking which way is the shortest path to wings level flight. The present study compared the baseline F-15E electronic attitude display to two other displays: the F-15E display with sky pointer arrows and a 180-degree field-of-view-dimensional display with sky pointer arrows. The results indicated significantly faster decision times for the baseline F-15E display (1.03 seconds) and F-15E display with arrows (1.00 seconds) relative to the three-dimensional display (1.18 seconds). This difference was present only for the 55-degree nose high attitude conditions. Nose low attitudes were significantly different in decision time. Correct responses using the F-15E display with arrows (90 percent) were significantly higher than the baseline F-15E display (76 percent) for nose low attitudes. GRA 07 AIRCRAFT PROPULSION AND POWER Includes prime propulsion systems and systems components, e.g., gas turbine engines and compressors; and onboard auxilliary power plants for aircraft. For related information see also 20 Spacecraft Propulsion and Power, 28 Propellants and Fuels, and 44 Energy Production and Conversion. N89-26847 Department of the Air Force, Washington, DC. William M. Madden, inventor (to AF), Claude R. Stogner, inventor (to AF), and Charles E. Spaeth, inventor (to AF) 12 Jul. 1988 4 p Filed 24 Feb. 1987 (AD-D014002; US-Patent-4,756,053; US-Patent-Appl-SN-018118; US-Patent-Class-162-23) Avail: US Patent and Trademark Office CSCL 13/5 This patent relates to a hinge joint for a nozzle flap for a gas turbine engine exhaust nozzle. First and second flaps are joined at a hinge joint formed by a hollow hinge pin passing through alternating lug extensions of the first and second flaps. A plurality of flow openings in the hollow hinge pin are provided for establishing fluid communication between the interior of the first flap and the interior of the second flap. N89-26848 Department of the Air Force, Washington, DC. GRA Walter D. Hutto, Jr., inventor (to AF) and William W. Stockton, inventor (to AF) 3 Jan. 1989 13 p Filed 14 May 1987 (AD-D014051; US-Patent-4,794,755; US-Patent-Appl-SN-049351; US-Patent-Class-60-39-281) Avail: US Patent and Trademark CSCL 01/4 Office A back-up control system is implemented in a single engine aircraft to provide inactivation of a faulty primary system and engagement of secondary system, and thereby provide a means of maintaining controllable flight sustaining thrust. The aircraft's hydromechanical main engine control and its companion pressure and temperature sensors can develop faults which can result in the inability of the engine to deliver flight sustaining thrust. The electronic control contains logic functions which indicate failure of the main engine control when all of the following exists for a minimum of three seconds: (1) the power lever is at a position requesting a level of dry thrust which exceeds a predetermined threshold; (2) core engine speed is below that required to deliver the predetermined level of dry thrust; (3) core engine speed is not increasing; and (4) turbine temperature is beneath the maximum allowable limit. The backup system is used in conjunction with a three position cockpit switch having normal, on and standby position. In the normal position the back-up system is off and must be manually activated by switching to the on position. In the standby position, the backup system is automatically activated when the necessary conditions occur. GRA N89-26849 Department of the Air Force, Washington, DC. VARIABLE NOZZLE AREA TURBINE VANE COOLING Patent Edward S. Hsia, inventor (to AF), John H. Starkweather, inventor 17 Jan. 1989 (to AF), and William K. Koffel, inventor (to AF) 13 p Filed 19 May 1986 (AD-D014071; US-Patent-4,798,515; US-Patent-Appl-SN-895016; US-Patent-Class-415-115) Avail: US Patent and Trademark Office CSCL 21/5 The cooling insert for a movable vane in a jet engine is divided into plural overlapping forward and aft segregated members, each member being coolant fed through separate vane trunnion areas AIRCRAFT PROPULSION AND POWER and each insert is of successively decreasing cross-sectional area. Problems of locally inadequate vane cooling, coolant match point movement, and unequal coolant supply pressures are addressed by the disclosed apparatus. Reuse of the trunnion supplied impingement cooling air for additional cooling functions is also disclosed. GRA N89-26850 Department of the Air Force, Washington, DC. Rosario N. Demers, inventor (to AF) 21 Feb. 1989 The system is used to measure the gap between the turbine engine compressor blade tip and the compressor case. The measurement is accomplished while the engine is running. This system has several features which minimize problems plaguing earlier systems. These include tuning stability and sensitivity drift. Both these problems are intensified by the environmental factors present on compressors, i.e., wide temperature fluctuations, vibrations, conductive contamination of probe tips and others. The circuitry in this new system provides phase lock feedback to control tuning and shunt calibration to measure sensitivity. The use of high frequency excitation lowers the probe tip impedance, thus minimizing the effects of contamination. The ability to control tuning and to calibrate has been demonstrated. N89-26851# GRA Naval Air Propulsion Test Center, Trenton, NJ. Propulsion Engineering Dept. STATISTICS ON AIRCRAFT GAS TURBINE ENGINE ROTOR FAILURES THAT OCCURRED IN US COMMERCIAL AVIATION DURING 1983 Final Report R. A. Delucia and J. T. Salvino Mar. 1989 (Contract DOT-FA71NAA-P98) 23 P (AD-A207592; NAPC-PE-184; DOT/FAA/CT-89/5) Avail: NTIS HC A03/MF A01 CSCL 21/5 This report presents statistics relating to gas turbine engine rotor failures which occurred during 1983 in commercial aviation service use. One-hundred and seventy-two failures occurred in 1983. Rotor fragments were generated in 96 of the failures and, of these, 9 were uncontained. The predominant failures involved blade fragments, 95.4 percent of which were contained. Five disk failures occurred and four were uncontained. Fifty-nine percent of the 172 failures occurred during the takeoff and climb stages of flight. This service data analysis is prepared on a calendar year basis and published yearly. The data support flight safety analyses, proposed regulatory actions, certification standards, and cost GRA benefit analyses. N89-26852 Department of the Air Force, Washington, DC. AFTERBURNER FLAMEHOLDER CONSTRUCTION Patent Donald W. Eldredge, inventor (to AF) and Billy R. Milam, inventor (to AF) 28 Mar. 1989 4 p Filed 25 Jun. 1987 (AD-D014116; US-Patent-4,815,283; US-Patent-Appl-SN-066154; US-Patent-Class-602-610) Avail: US Patent and Trademark Office CSCL 21/5 A flameholder for a gas turbine engine includes a retainer plate on the igniter holder boss which protrudes from the flameholder gutter and extends through a clearance hole in the outer shroud. The retainer plate permits thermally induced growth of the parts but does not allow the fuel spray and air flow gap near the igniter to increase significantly, thus assuredly maintaining GRA a sufficiently rich fuel/air mixture to promote ignition. N89-26853 Department of the Air Force, Washington, DC. Harold R. Hansel, inventor (to AF) and Vincent M. Drerup, inventor (to AF) 28 Mar. 1989 5 p Filed 13 Nov. 1987 (AD-D014123; US-Patent-4,815,933; US-Patent-Appl-SN-122153; N89-26854# Advisory Group for Aerospace Research and Development, Neuilly-Sur-Seine (France). MEASUREMENT UNCERTAINTY WITHIN THE UNIFORM J. P. K. Vleghert, ed. (National Aerospace Lab., Amsterdam, (AGARD-AG-307; ISBN-92-835-0508-5) Copyright Avail: NTIS HC A05/MF A01 This AGARDograph is an outcome of the Propulsion and Energetics Panel Working Group 15 on, Uniform Engine Testing Programme (AGARD AR 248). During the performance of this Group it appeared that the results of some test runs were somewhat scattered, without an obvious explanation. The Group, therefore, formed a sub-Group with the task of carefully assessing the uncertainties of the measured data in order to find out whether the scattering was within the expected uncertainty or whether an explanation must be found. Since the results of the efforts of the sub-Group have some importance beyond the Working Group 15 tests, it was decided to report them in the form of an AGARDograph. In Chapter 5 the different uncertainties are estimated. The discussion on the uncertainties appears in Chapter 6 and in the following Chapter 7, ten conclusions are drawn from the efforts. This AGARDograph was prepared at the request of the Propulsion and Energetics Panel of AGARD. Author N89-26855# Sandia National Labs., Albuquerque, NM. Advanced Technology Div. AUTONOMOUS LAND NAVIGATION: A DEMONSTRATION OF RETROTRAVERSE P. R. Klarer 1989 8 p Presented at the 16th AUVS Annual Technical Symposium and Exhibit, Washington, DC, 17 Jul. 1989 (Contract DE-AC04-76DP-00789) (DE89-011881; SAND-89-0393C; CONF-890783-1) Avail: NTIS HC A02/MF A01 This paper describes a hardware and software system developed to perform autonomous navigation of a land vehicle in a structured environment. The vehicle used for development and testing of the system was the Jeep Cherokee Mobile Robotics Testbed Vehicle developed at Sandia National Laboratories in Albuquerque. Since obstacle detection/avoidance has not yet been incorporated into the system, a structured environment is postulated that presumes the paths to be traversed are obstacle-free. The system performs path planning and execution (following) based on maps constructed using the vehicle's navigation system and onboard map-maker. The system configuration allows a map to be generated and stored during teleoperation of the vehicle, which may then be inverted and autonomously followed to perform retrotraverse back to the path start point. The system software, hardware, and performance data are discussed. DOE N89-26856# Astron Research and Engineering, Sunnyvale, CA. DEMONSTRATION OF OBLIQUE DETONATION WAVE FOR HYPERSONIC PROPULSION Final Technical Report, 1 Aug. 1988-31 Jan. 1989 Takashi Nakamura, Michael J. Schuh, Donald S. Randall, Thomas J. Dahm, and David T. Pratt 30 Mar. 1989 (Contract F49620-88-C-0130) 118 P (AD-A208268; ASTRON-7151-001; AFOSR-89-0659TR) Avail: NTIS HC A06/MF A01 CSCL 21/5 The Oblique Detonation Wave Engine (ODWE) offers a number of advantages over the Supersonic Combustor Ramjet (SCRAMJET) for hypersonic aeropropulsion. The objective of this program is to obtain data on the stability of the Oblique Detonation Wave (ODW) and to assess the applicability of the ODW to hypersonic propulsion. The program consists of the basic study of the ODW phenomenon and the design study of the test facility (Phase 1), and an indepth experimental study of the ODW in a ram cannon-type combustion tube with a hypervelocity projectile launched into the tube by a two-stage light-gas gun (Phase 2). This Phase 1 report summarizes the results pertaining to the stability of the ODW and the experimental facility designs. It is concluded that ODW will be initiated and sustained in the test facility configuration and that the tests will generate data concerning key issues for the application of the ODW to hypersonic propulsion. 08 AIRCRAFT STABILITY AND CONTROL GRA Includes aircraft handling qualities; piloting; flight controls; and autopilots. For related information see also 05 Aircraft Design, Testing and Performance. N89-26857 Department of the Air Force, Washington, DC. CONTROL SURFACE DUAL REDUNDANT SERVOMECHANISM Patent Clete M. Boldrin, inventor (to AF), Richard D. McCorkle, inventor (to AF), Jimmy W. Rice, inventor (to AF), and James J. Rustik, inventor (to AF) 31 Jan. 1989 6 p Filed 11 Dec. 1984 (AD-D014058; US-Patent-4,800,798; US-Patent-Appl-SN-680674) Avail: US Patent and Trademark Office CSCL 09/1 A dual redundant servomechanism for moving aircraft control surfaces is disclosed. The servomechanism is of the type whose input commands, from the pilot of the aircraft, are transmitted electrically. Force fight, which is associated with such dual servomechanisms when they are connected to a common aircraft control surface, is minimized. This is accomplished by providing the control system for each servomechanism with input signals which are electrically summed. Each control system includes electrical transducers which provide a signal indicative of actuator position and the pressure associated with the hydraulic motor used in each servomechanism. GRA N89-26858 Department of the Air Force, Washington, DC. BRAIN 02 RESERVE LIMITER FOR HIGH PERFORMANCE AIRCRAFT Patent 18 Apr. 1989 6 p Robert E. Van Patten, inventor (to AF) A method for controlling an aircraft to prevent high G-caused pilot unconsciousness is disclosed. Data defining a state space of acceleration, rate of change of acceleration and duration of acceleration at maximum acceleration within which an aircraft may be operated without causing pilot unconsciousness is provided to an aircraft intelligent flight control system. The flight control system continuously monitors the past and present state of the aircraft and compares to the surface boundaries of the defined safe state space. Whenever the aircraft exceeds those boundaries, the flight control system intervenes to unload the aircraft to within those boundaries. Additional data and measurements may be added to define an n-dimensional state space. Another embodiment unloads the aircraft to a baseline acceleration. A simplified embodiment is described which compares current acceleration to a preselected value of acceleration. If the current acceleration exceeds the preselected value, the previous acceleration onset rate is compared to a preselected acceleration onset rate. If it exceeds the preselected onset rate, the duration of time the current and immediately past acceleration has exceeded the preselected value of acceleration is determined. If that value is greater than a preselected duration, the flight control system commands the aircraft ot perform an unloading maneuver to reduce the G loading on the aircraft to a preselected baseline acceleration. GRA Princeton Univ., NJ. Dept. of Mechanical and Aerospace Engineering. N89-26859*# DESIGN OF AN ACTIVE HELICOPTER CONTROL (NASA-CR-185490; NAS 1.26:185490) Avail: NTIS HC In an effort to develop an active control technique for reducing helicopter vibrations stemming from the main rotor system, a helicopter model was designed and tested at the Princeton Rotorcraft Dynamics Laboratory (PRDL). A description of this facility, including its latest data acquisition upgrade, are given. The design procedures for the test model and its Froude scaled rotor system are also discussed. The approach for performing active control is based on the idea that rotor states can be identified by instrumenting the rotor blades. Using this knowledge, Individual Blade Control (IBC) or Higher Harmonic Control (HHC) pitch input commands may be used to impact on rotor dynamics in such a way as to reduce rotor vibrations. Discussed here is an instrumentation configuration utilizing miniature accelerometers to measure and estimate first and second out-of-plane bending mode positions and velocities. To verify this technique, the model was tested, and resulting data were used to estimate rotor states as well as flap and bending coefficients, procedures for which are discussed. Overall results show that a cost- and time-effective method for building a useful test model for future active control experiments was developed. With some fine-tuning or slight adjustments in sensor configuration, prospects for obtaining good state estimates look promising. Author (NASA-CR-4111; H-1438; NAS 1.26:4111) Avail: NTIS HC A11/MF A01 CSCL 01C Feb. A fixed-base simulation was performed to identify and quantify interactions between the pilot's hand/arm neuromuscular subsystem and such control system features of typical modern fighter aircraft roll rate command mechanizations as: (1) force versus displacement sensing side-stick type manipulator, (2) feel force/displacement gradient, (3) feel system versus command prefilter dynamic lag, and (4) flight control system effective time delay. The experiment encompassed some 48 manipulator/ Displacement side-stick exfilter/aircraft configurations. periment results are given and compared with the previous force sidestick experiment results. Attention is focused on control bandwidth, excitement (peaking) of the neuromuscular mode, feel force/displacement gradient effects, time delay effects, etc. Section 5 is devoted to experiments with a center-stick in which force versus displacement sensing, feel system lag, and command prefilter lag influences on tracking performance and pilot preference are investigated. N89-26861*# Author National Aeronautics and Space Administration. Langley Research Center, Hampton, VA. AERODYNAMIC PARAMETERS OF AN ADVANCED FIGHTER AIRCRAFT ESTIMATED FROM FLIGHT DATA. PRELIMINARY RESULTS Vladislav Klein, Keven P. Breneman, and Thomas P. Ratvasky (Joint Inst. for Advancement of Flight Sciences, Hampton, VA.) Jul. 1989 59 p (NASA-TM-101631; NAS 1.15:101631) Avail: NTIS HC A01 Results of a predevelopment program for Fly by Wire (FBW) flight control systems in civil aircraft are presented. The concept is based on requirements for certification of safety-critical functions, capacity of the computer systems to implement active control function (gust load alleviation, variable camber wing control etc) and cockpit standard with mini-stick and digital display and operation systems. The experimental system consists of a quad-redundant, fault-tolerant multicomputer system and a quadruplex Actuator Electronic Unit (AEU) providing the redundancy management and force synchronization of parallel-active electrohydraulic servoactuators. The computer-interlane communication is by fiber optics. By A300/310 flight simulator-investigations the control laws with command and flight envelope protections are validated. The experimental system was tested in a rig with aircraft-simulation in the loop. The control and redundancy management-concept for the AEU's and parallel-redundant actuator configurations are verified ESA by test results. N89-26863# Naval Postgraduate School, Monterey, CA. (AD-A208461) Avail: NTIS HC A06/MF A01 CSCL 01/4 New mission requirements dictate the need to improve the P-3's defensive maneuvering capabilities. Research was conducted to find viable methods of increasing the current roll response of the P-3. First, a flight simulator was used to determine an initial target roll response. Next, a computer code was used to evaluate the aerodynamic effect of varying the size and deflection of the aileron. These results, along with the flight simulator tests, were used to analyze the requirements to reach the target response. Several ways to achieve this goal are discussed. It was found that by increasing the aileron deflection from + or - 20 deg to + or - 25 deg and increasing the aileron chord by 50 percent, a 58 percent increase in Ci (total rolling moment coefficient) could be realized. This does not reach the goal of a 100 percent increase in Ci, but, it does yield a large increase in lateral control response. An increase in aileron size and deflection along with some of the other suggested modifications would certainly approach the desired GRA goal. Tony B. Husbands and Dennis L. Bean Apr. 1989 93 p (AD-A207554; CTIAC-84; WES/TR/SL-87-33; WES/TR/SL-89-5) Avail: NTIS HC A05/MF A01 CSCL 01/4 The U.S. Air Force uses an aircraft arresting system on many of their runways for emergency stopping of aircraft. It consists of a 1 or 1-1/4-inch steel cable stretched across the runway connected to a braking mechanism. When aircraft tires impact the cable, considerable damage occurs to concrete and other materials placed underneath the cable. Materials previously used were not performing satisfactorily for various reasons. A survey was made in 1980 for the Air Force Engineering and Services Center (AFESC) to locate materials for evaluation. Five of these were selected for detailed testing. The materials were tested for gel times, peak exotherms, bond strength, abrasion resistivity, ultraviolet degradation, resiliency, hardness, abrasion-impact resistance, effect of curing temperature, and proportioning errors. The materials which showed most promise were field tested at Homestead and Tyndall Air Force Bases. GRA A03/MF A01 CSCL 14B Real time analysis of data can reduce the time involved in exploring dynamic systems. The failure of the data acquisition system at the Princeton Dynamic Model Track prompted its replacement with a real time data acquisition system. Data can be obtained from an experiment and analyzed during and immediately following a data run. The new system employs high speed analog to digital conversion and a small computer to collect data. Sampling rates of 1000 hertz over 44 channels (44,000 words/sec) are obtainable. The data can be accessed as it enters the computer's environment where it may be displayed or stored for later processing. The system was tested on a helicopter rotor steep descent experiment. The data collected compares with previous data from a similar experiment. Author Stanford Univ., CA. Dept. of Aeronautics and DESIGN AND CALIBRATION OF THE MIXING LAYER AND James H. Bell and Rabindra D. Mehta May 1989 (NASA-CR-185472; NAS 1.26:185472; JIAA-TR-89) HC A03/MF A01 CSCL 01E 36 P Avail: NTIS A detailed account of the design, assembly and calibration of a wind tunnel specifically designed for free-shear layer research is contained. The construction of this new facility was motivated by a strong interest in the study of plane mixing layers with varying initial and operating conditions. The Mixing Layer Wind tunnel is located in the Fluid Mechanics Laboratory at NASA Ames Research Center. The tunnel consists of two separate legs which are driven independently by centrifugal blowers connected to variable speed motors. The blower/motor combinations are sized such that one is smaller than the other, giving maximum flow speeds of about 20 and 40 m/s, respectively. The blower speeds can either be set manually or via the Microvax II computer. The two streams are allowed to merge in the test section at the sharp trailing edge of a slowly tapering splitter plate. The test section is 36 cm in the cross-stream direction, 91 cm in the spanwise direction and 366 cm in length. One test section side-wall is slotted for probe access and adjustable so that the streamwise pressure gradient may be controlled. The wind tunnel is also equipped with a computer controlled, three-dimensional traversing system which is used to flow fields with investigate the and hot-wire pressure instrumentation. The wind tunnel calibration results show that the mean flow in the test section is uniform to within plus or minus 0.25 pct and the flow angularity is less than 0.25 deg. The total streamwise free-stream turbulence intensity level is approximately 0.15 pct. Currently the wind tunnel is being used in experiments designed to study the three-dimensional structure of plane mixing layers and wakes. Author N89-26867# Wisconsin Univ., Madison. Applied Superconductivity DESIGN AND OPERATION OF A HORIZONTAL LIQUID S. W. VanSciver and J. G. Wiesend, II 1988 6 p Presented at the 12th International Cryogenic Engineering Conference and Exhibition, Southampton, United Kingdom, 12-15 Jul. 1988 (Contract DE-AC02-86ER-40306) (DE89-013482; CONF-880736-6) Avail: NTIS HC A02/MF A01 The University of Wisconsin horizontal liquid helium flow facility (LHFF) consists of a five meter long 20 cm ID horizontal dewar connected to two end boxes. Several heat exchanger inserts have been built to allow variable temperature operation of 1.6 K less than or equal to T less than or equal to 4.2 K. A centrifugal pump is installed at one end of the facility permitting experiments in forced flow liquid helium up to 100 gm/s. The horizontal design allows experimentation on long straight test sections which may be used either to study fundamental properties of heat and mass transfer in helium or prototype cryogenic components under realistic conditions. A detailed description of the design and operating experience of the LHFF is presented. DOE N89-26868# Federal Aviation Administration, Atlantic City, NJ. (DOT/FAA/CT-TN89/45) Avail: NTIS HC A03/MF A01 An updated version of an 18-inch internally lighted wind cone was evaluated. The original model was evaluated at the Federal Aviation Administration (FAA) Technical Center. The results of that evaluation, as detailed in Technical Note DOT/FAA/CT-TN85/4 (February 1985), revealed that the original model of the internally lighted wind cone did not provide adequate wind direction and speed information under low velocity (10 knots or less) wind conditions. The new model was installed in close proximity to the standard externally lighted wind cone at the Atlantic City International Airport, N.J., to permit a comparative evaluation by FAA and general aviation pilots. Results of the testing, accomplished during taxi and flight operations under wind conditions of 10 knots or less, revealed that the updated model still does not provide adequate wind direction and speed information. Participating pilots preferred the standard externally lighted wind cone. Author |