SIMSCRIPT. Events in the simulation included takeoffs. departures, enroutes, missions, arrivals at Initial Approach Fixes (IAFs) and landings. To demonstrate potential use of the model, the problem of rescheduling Strategic Air Command (SAC) aircraft upon base closures was addressed. Two strategies for the diversion of such aircraft were developed, implemented and the results compared on the basis of impact on final destination bases and average aircraft airborne time. Strategy 1 entailed the rerouting of aircraft to designated alternate bases or to the nearest open base without constraint. Strategy 2 involved the selection of an alternate base by insuring that parking spaces and appropriate maintenance support were available. Finally, extensions to the model and recommendations are discussed. GRA N79-31190# National Technical Information Service, Springfield, Va. LORAN C NAVIGATION. CITATIONS FROM THE NTIS (NTIS/PS-79/0523/5; NTIS/PS-78/0531; NTIS/PS-77/0336) Avail: NTIS HC $28.00/MF $28.00 CSCL 17G Research reports are cited on the use of Loran C for air and maritime navigation, including coastal confluence and inland waters; automatic vehicle location and tracking; distress systems; portable land position finding: meteorological and atmospheric data acquisition; time and frequency synchronization; and position fixing. Also included are research on Loran C equipment and signal propagation. This updated bibliography contains 144 abstracts, 21 of which are new entries to the previous edition. GRA N79-31191# National Technical Information Service, Springfield, Va. LORAN C NAVIGATION. CITATIONS FROM THE ENGINEERING INDEX DATA BASE Progress Report, 1970 May 1979 William E. Reed Jun. 1979 130 p Supersedes NTIS/PS-78/ 0532; NTIS/PS-77/0337 (NTIS/PS-79/0524/3; NTIS/PS-78/0532; NTIS/PS-77/0337) Avail: NTIS HC $28.00/MF $28.00 CSCL 17G The use of Loran C for air and maritime navigation is discussed Topics covered include both automatic vehicle locating and tracking; meteorological and atmospheric data acquisition; position fixing, and time and frequency synchronization as well as Loran C equipment and signal propagation. This updated bibliography contains 124 abstracts, 15 of which are new entries to the previous edition. GRA Pilot performance during a terrain following flight was studied for ride quality criteria validation. Data from manual and automatic terrain following operations conducted during low level penetrations were analyzed to determine the effect of ride qualities on crew performance. The conditions analyzed included varying levels of turbulence, terrain roughness, and mission duration with a ride smoothing system on and off. Limited validation of the B-1 ride quality criteria and some of the first order interactions between ride qualities and pilot/vehicle performance are highlighted. An earlier B-1 flight simulation program correlated well with the flight test results. A.W.H. N79-31195*# Lockheed-Georgia Co., Marietta. J. F. Cahill and P. C. Connor Washington NASA Sep. 1979 82 p refs (Contract NAS2-9331) (NASA-CR-3178) Avail: NTIS HC A05/MF A01 CSCL 01C Pressure data from a number of previous wind tunnel and flight investigations of high speed transport type wings were analyzed with the intent of developing a procedure for extrapolating low Reynolds number data to flight conditions. These analyses produced a correlation of the development of trailing-edge separation resulting from increases in Mach number and/or angle of attack and show that scale effects on this correlated separation development and the resulting shock location changes fall into a regular and apparently universal pattern. Further studies appear warranted to refine the correlation through a detailed consideration of boundary layer characteristics, and to evaluate scale effects on supercritical wings. Author N79-31196# Naval Postgraduate School, Monterey, Calif. Wayne Richard Hanley Mar. 1979 62 p refs This thesis examines the failure at a fastener hole in a composite fuel tank skin due to hydraulic ram i.e., fluid pressure due to a penetrating projectile. An experimental apparatus was set up to investigate the triaxial loading conditions at the fastener hole so that an M-P-N failure surface could be developed for the graphite-epoxy laminate. An expression was derived to predict the bending moment at the fastener in terms of the pull force on the fastener and the axial force in the plate. Aluminum specimens were tested and the results were compared with the predicted results to validate the experimental procedure. The experimental results were found to not be repeatable, and hence a correlation with the predicted results was not appropriate. The nonrepeatability could not be explained. Composite specimens were fabricated and prepared for testing Experimental values for the ultimate pure pull force. P. ultimate pure bending moment, M, and ultimate pure axial force, N, were obtained. This program investigated analytically and experimentally the effect of transport/bomber loads spectrum variations on crack growth The spectrum represented a STOL transport (C-15) wing lower surface loading The 116 spectrum variations were generated, grouped in the following categories (1) baseline spectra, (2) mission mix, (3) sequence of missions, (4) individual flight length, (5) flight segments, (6) exceedances spectra, (7) design stress level, (8) valley/peak coupling, (9) low load truncation, (10) high infrequent loads, (11) clipping of large loads. (12) miscellaneous variations, and (13) combined variations. Spectra were generated as random cycle-by-cycle, flight-by-flight sequences. Analyses and tests were performed on 7475-T7651 aluminum, to 25 inches, starting with an initial through-thethickness 0.03-inch crack out of a 1/4-inch diameter hole Crack growth analysis predictions were made for all 116 spectra using the linear model. Crack growth tests were performed with 33 of these spectra. Good correlation was obtained between analysis and test results in all cases except with spectra dealing with increased frequency and magnitude of high infrequent loads and spectra which were drastically changed from a wing-type to a vertical tail-type spectrum. Largest effects on crack growth life, as measured in flight hours, was due to flight length, mission, mission mix, and design stress level variations. Based on the results of this program, fleetwide crack growth variations by a factor of 100 and 10 could be experienced, depending on whether it was short-term or long-term variation. GRA The report contains, in the form of a user's manual, the listing and complete description of a computer program to generate fatigue spectrum loading sequences. The program is specifically tailored for the development of random cycle-by-cycle, flight-byflight loading sequences typical of aircraft structures. However, its general features allow the development of any type spectrum. The random sequence of cycles and flights is produced by a random number generator. Alternate non-random flight sequences can also be generated. The basic input data consists of loads exceedances spectra or data to calculate such spectra by the program. The program contains the following spectrum editing features: (1) truncation elimination of cycles as a function of range and R, peak or valley, (2) clipping loads below or above a specified clipping value are set equal to that value, (3) all loads in the spectrum are multiplied by a constant. The output is a valley, peak sequence of the loads spectrum. The spectrum may be read into a magnetic tape to be used in other analyses or in testing. GRA N79-31199# Air Force Inst. of Tech., Wright-Patterson AFB, Ohio School of Engineering. INVESTIGATION OF ROLL PERFORMANCE FOR A HIGHLY A linearized model of a fighter-type aircraft with significant roll-pitch inertial coupling, including its full flight control system is required in order to conduct a comparative analysis between body axis rolls and velocity (stability) axis rolls. The stability of the aircraft is checked at various roll rates for both body axis and velocity axis rolls. This is done by examining the signs of the eigenvalues of the linearized model-positive for unstable and negative for stable. It is found that at various angles of attack the velocity axis rolls prove to be at least as stable and, in most cases, more stable than body axis rolls. The stability is also observable for various combinations of flight control systems. In developing a nonlinear coupled equation solver, a single equation with known solutions is considered first. This is done to show a simplified version of what the nonlinear program is required to do. Next, a pair of nonlinear coupled equations is analyzed. The development of the program for the single equation case proves to be successful, but certain problems arise when working with a pair of coupled equations. This thesis provides a good foothold on a method of analysis known as Bifurcation Analysis and Catastrophe Theory which can be used to solve the nonlinear coupled aircraft equations. This thesis presents some GRA of the problems which could be encountered This report documents the work accomplished for the Windshield Technology Demonstrator Program The studies. analyses, testing and development accomplished during this program involved a total system approach required for aircraft canopies in the context of the continuing Air Force generic wind shield development programs. State-of-the-art applications of new transparency materials have been devised from both military and commercial aircraft with major attention directed to the topics of bird impact resistance, structural design integration, systems integration, and design for maintainability and reliability The authors and Air Force Flight Dynamics Laboratory FEW agree that the various disciplines and essential technical design concepts represented, including associated reports noted in Section IV. should all be utilized in arriving at an optimum design of the canopy system for any production aircraft. GRA N79-31202# Dayton Univ., Ohio. Research Inst. Failure data, displacement data, and pressure data were obtained from laboratory experiments. Panels were made from 7075-T6 and 2024- T3 aluminum and from graphite epoxy panel thicknesses were 1.6 to 6.35 mm. Protection included 10 mm ballistic foam and stiffeners. Projectiles were 5.6 g and 11.7 g spheres and cubes. Failures were always catastrophic, and failure thresholds were always abrupt. When cracks formed, they ran across the panels, except when stiffeners were present. In thin panels, cracks initiated at the corners of the perforation when cubical fragments were used. The entrance panel damage was primarily induced by the shock wave generated by the impact. The very high shock pressure resulted in impulsive loading of the panels that caused prompt crack formation. Cracks were propagated by the displacement field. GRA Several aero-acoustic suppression devices have been evaluated which were considered feasible for installation on an F-111 aircraft for flight test evaluation. The most promising modification consists of a saw tooth spoiler mounted at the leading edge of the weapon bay. This device would be erected to a 90 degree position during the bay doors opening sequence. The spoiler is folded flush with the fuselage during all other flight conditions. Wind tunnel tests have shown that this spoiler improves the aero-acoustic environment within the open weapon bay and improves the weapon separation characteristics over the Mach range of 95 to 1.3 investigated during the drop test phase. GRA N79-31204# Argonne National Lab., III. ANALYSIS OF PLUME RISE FROM JET AIRCRAFT R. J. Yamartino, J. Lee, S. Bremer, D. Smith, and J. Calman 1979 7 p refs Presented at the 4th Symp. on Turbulence, Diffusion and Air Pollution, Boston, 15-18 Jan. 1979 Sponsored in part by FAA and Air Force Prepared in cooperation with Environmental Research and Technology, Lexington, Mass (Contract W-31-109-eng-38) (CONF-790142-1) Avail: NTIS HC A02/MF A01 Preliminary results of an investigation of the behavior of buoyant jet engine exhaust plumes in a crosswind are discussed and attempts are made to identify the degree to which the plume rise can be described by relationships developed for other types of sources. At least four factors were found to affect the rate of dilution of jet exhaust before it reached receptors adjacent to taxiways or runways: (1) turbulent mixing of the jet exhaust at the engine exit; (2) buoyant plume rise; (3) advective dilution; and (4) dispersion by ambient turbulence. N79-31701# Observatoire de Paris-Meudon (France). SPACE RESEARCH WITH AIRBORNE PLATFORMS DOE P. Lena and D. Rouan In ESA European Sounding Rocket, Balloon and Related Res., with Emphasis on Expt. at High Latitudes Jun. 1978 p 461-465 refs (For primary document see N79-31637 22-42) Avail: NTIS HC A23/MF A01 The Advanced Systems Division of the U.S. Army Avionics Research and Development Activity at Fort Monmouth, New Jersey, has issued Engineering Development (ED) contracts to two contractors for the Integrated Avionics Control System (IACS). One of the features of the ED progam will be a complete logistics support analysis (LSA) of Reliability Improvement Warranty (RIW) as an alternative to Army organic support. ARINC Research Corportion assisted the IACS Project Office in the development of draft RIW terms and conditions on which the LSA will be based. This report presents the activities that were performed and describes the draft RIW and conditions that were developed. GRA Jets flowing from air entry holes of the combustor liner of a gas turbine were investigated. Cold air was supplied through the air entry holes into the primary hot gas flows. The mass flow of the primary hot gas and issuing jets was measured, and the behavior of the air jets was studied by the measurement of the temperature distribution of the gas mixture. The air jets flowing from three circular air entry holes, single streamwise long holes, and two opposing circular holes, parallel to the primary flow were studied along with the effects of jet and gas stream velocities, and of gas temperature. The discharge coefficient, the maximum penetration of the jets, the jet flow path, the mixing of the jets, and temperature distribution across the jets were investigated. Empirical expressions which describe the characteristics of the jets under the conditions of the experiments were formulated. A.W.H. pollutant emission levels was evaluated in a series of CF6-50 engine tests. Engine lightoff was readily obtained and no difficulties were encountered with combustor staging. Engine acceleration and deceleration were smooth, responsive and essentially the same as those obtainable with the CF6-50 combustor. The emission reductions obtained in carbon monoxide, hydrocarbons, and nitrogen oxide levels were 55, 95, and 30 percent. respectively, at an idle power setting of 3.3 percent of takeoff power on an EPA parameter basis. Acceptable smoke levels were also obtained. The exit temperature distribution of the combustor was found to be its major performance deficiency. In all other important combustion system performance aspects, the combustor was found to be generally satisfactory. K.L. The JT9D-70/59 high pressure turbine active clearance control system was modified to provide reduction of blade tip clearance when the system is activated during cruise operation. The modification increased the flow capacity and air impingement effectiveness of the cooling air manifold to augment turbine case shrinkage capability, and increased responsiveness of the airseal clearance to case shrinkage. The simulated altitude engine testing indicated a significant improvement in specific fuel consumption with the modified system. A 1000 cycle engine endurance test showed no unusual wear or performance deterioration effects on the engine or the clearance control system. Rig tests indicated that the air impingement and seal support configurations used in the engine tests are near optimum. KL. The TSIO-360-C engine (S/N 300244) was tested to develop an exhaust emissions data base. This data base consists of currents of current production baseline emissions characteristics, lean-out emissions data, effects of leaning-out the fuel schedule on cylinder head temperatures, and data showing ambient effects on exhaust emissions and cylinder head temperatures. The engine operating with its current full-rich production fuel schedule could not meet the proposed Environmental Protection Agency (EPA) standard for carbon monoxide (CO) and unburned hydrocarbons (HC) under sea level standard-day conditions. The engine did, however, meet the proposed EPA standard for oxides of nitrogen (NOX) under the same sea level conditions. The results of engine testing under different ambient conditions (essentially sea level standard day to sea level hot day) are also presented, and these results show a trend toward higher levels of emissions output for CO and HC while producing slightly lower levels of NOX. J.M.S. N79-31212*# United Technologies Corp., East Hartford, Conn. AERODYNAMIC AND ACOUSTIC INVESTIGATION OF INVERTED VELOCITY PROFILE COANNULAR EXHAUST NOZZLE MODELS AND DEVELOPMENT OF AERODYNAMIC AND ACOUSTIC PREDICTION PROCEDURES Final Report Richard S. Larson, Douglas P. Nelson, and Bradley S. Stevens Washington NASA Aug. 1979 223 p refs (Contract NAS3-20061) (NASA-CR-3168: PWA-5550-8) HC A10/MF A01 CSCL 21E Avail: NTIS Five co-annular nozzle models, covering a systematic variation of nozzle geometry, were tested statically over a range of exhaust conditions including inverted velocity profile (IVP) (fan to primary stream velocity ratio > 1) and non IVP profiles. Fan nozzle pressure ratio (FNPR) was varied from 1.3 to 4.1 at primary nozzle pressure ratios (PNPR) of 1.53 and 2.0. Fan stream temperatures of 700 K (1260 deg R) and 1089 K(1960 deg R) were tested with primary stream temperatures of 700 K (1260 deg R), 811 K (1460 deg R), and 1089 K (1960 deg R). At fan and primary stream velocities of 610 and 427 m/sec (2000 and 1400 ft/sec), respectively, increasing fan radius ratio from 0.69 to 0.83 reduced peak perceived noise level (PNL) 3 dB, and an increase in primary radius ratio from 0 to 0.81 (fan radius ratio constant at 0.83) reduced peak PNL an additional 1.0 dB. There were no noise reductions at a fan stream velocity of 853 m/sec (2800 ft/sec). Increasing fan radius ratio from 0.69 to 0.83 reduced nozzle thrust coefficient 1.2 to 1.5% at a PNPR of 1.53, and 1.7 to 2.0% at a PNPR of 2.0. The developed acoustic prediction procedure collapsed the existing data with standard deviation varying from or 8 dB to + or - 7 dB. The aerodynamic performance prediction procedure collapsed thrust coefficient measurements to within + or - 004 at a Author FNPR of 4.0 and a PNPR of 2.0. N79-31213*# National Aeronautics and Space Administration. Royce D. Moore and Walter M. Osborn Sep. 1979 71 p 21E The performance of a variable pitch fan stage tested over a range of blade setting angles, speeds, and flows is presented. The fan was designed for a tip speed of 289.6 m/sec and a flow of 29.6 kg/sec. The measured performance agreed reasonably well with the design point. The stall margin was only 5 percent. Static thrust values along an operating line ranged from less than 15 to over 115 percent of that at design angle as the blade setting angle was varied from 25 degrees (closed) to -8 degrees (opened). The use of casing treatment increased the stall margin to 20.6 percent but decreased efficiency by 4 percentage points. A.W.H. N79-31214*# National Aeronautics and Space Administration. Lewis Research Center. Cleveland, Ohio. AERODYNAMIC PERFORMANCE OF AXIAL-FLOW FAN STAGE OPERATED AT NINE INLET GUIDE VANE ANGLES Royce D. Moore and Lonnie Reid Sep. 1979 43 prefs (NASA-TP-1510; E-9714) Avail: NTIS HC A03/MF A01 CSCL 21E The overall performance of a fan stage with nine inlet guide vane angle settings is presented. These data were obtained over the stable flow range at speeds from 60 to 120 percent of design for vane setting angles from -25 to 42.5 degrees. At design speed and design inlet guide vane angle, the stage has a peak efficiency of 0.892 at a pressure ratio of 1.322 and a flow of 25.31 kg/s. The stall margin based on peak efficiency and stall was 20 percent. Based on an operating line passing through the peak efficiency point at the design setting angle. the useful operating range of the stage at design speed is limited by stall at the positive setting angles and by choke at the negative angles. At design the calculated static thrust along the operating line varied from 68 to 114 percent of that obtained at design' setting angle. Author N79-31215*# National Aeronautics and Space Administration. An inlet concept for separate turbojet and ramjet engines was defined and compared with an equivalent inlet for a wraparound turboramjet engine. The comparison was made for a typical high altitude hypersonic cruise vehicle where the turbojet inlet capture area was required to be half as large as the ramjet inlet capture area at cruise. The use of a shorter nacelle having substantially lower cooling requirements at cruise for the inlet concept for separate turbojet and ramjet engines is suggested. The separate engine concept better isolates the turbojet from the ramjet, requires no special close off mechanisms within the turbojet, and avoids the circumferential heat load imposed by a wraparound ramjet. A more variable geometry is required. A.W.H. N79-31216# North Carolina State Univ. at Raleigh. Engineering Design Center AN ACTUATOR DISK ANALYSIS OF AN ISOLATED ROTOR WITH DISTORTED INFLOW Progress Report, Jan. 1976 Dec. 1978 John N. Perkins 31 Mar. 1979 31 p Prepared in cooperation with United Technologies Research Center, E. Hartford, Conn. (Contract F44620-76-C-0055) (AD-A069884; NCSU/EDC-79-1) HC A03/MF A01 CSCL 21/5 Avail: NTIS An analytical study of the passage of distorted flow through an isolated, high hub-tip ratio axial compressor is reported. The analysis involves the coupling of the unsteady blade row aerodynamic response (with the flow immediately upstream of the blade row prescribed) with the unsteady duct flow both upstream and downstream of the blade row under the influence of the blade row loading. The numerical solution uses a starting procedure which does not require the inlet plane to be far enough upstream to be unaffected by the presence of the blade row. Hence any experimentally determined distortion at any arbitrary distance upstream of the blade row can be modelled. The results obtained indicate that the predicted pressure profile at the blade row exit is strongly dependent on the experimentally determined steady-state loss and exit flow angle curves, but is almost independent of the magnitude of the first order lag coefficient used to represent the boundary layer time delay through the blade passages. GRA for a new monitoring system--the Engine Diagnostic System-under development for the F100 engine on the F-15 and F-16 tactical fighter aircraft. The examination reveals that two different approaches to engine monitoring have evolved in attempts to achieve the goal of improved engine operations, maintenance. and management while reducing support costs. The first concentrates on short-term operations and maintenance aspects and is usually accomplished by recording inflight data in a snapshot mode, i.e., a few seconds of data either at predefined performance windows or when certain engine operating limits are exceeded. The second approach focuses on long-term design-oriented benefits through improved knowledge of the engine operating environment. To achieve the design-oriented benefits, data must be recorded continuously on at least a few aircraft at each operational location. GRA Recent investigations have indicated that aircraft engine exhaust emissions are sensitive to ambient conditions. This paper reports on combustor rig testing intended to evaluate variations due to ambient temperature and pressure with special emphasis on idle engine operating conditions. Empirically determined CO. C sub x, H sub y, and NO sub x correction factors--the ratio of the pollutant emission index value obtained during standard day operation to that resulting during actual ambient conditions--are presented. The effects of engine idle cycle pressure ratio, primary zone fuel-air ratio, and fuel type were investigated. Ambient temperature variations were seen to cause substantial emission changes; correction factors in excess of 2.0 were determined in some cases. Ambient pressure variations were found to be less substantial. A previously published NO sub x emission model and a simplified hydrocarbon combustion analysis are shown to be in general agreement with the empirical results. GRA N79-31219# General Electric Co., Lynn, Mass. Aircraft Engine Group. DEVELOPMENT OF HOT ISOSTATICALLY PRESSED RENE 95 TURBINE PARTS, ADDENDUM, PHASE 2 Final Report, 1976 - 1978 P. S. Mathur and J. L. Bartos Apr. 1979 44 p ref Addendum to USAAMRDL-TR-76-30 (Contract DAAJ02-73-C-0106) (AD-A069979; USARTL-TR-78-56-Add: USAAMRDL-TR-76-30) Avail: NTIS HC A03/MF A01 CSCL 21/5 This report deals with the results of Task 6 Cooling Rate Analysis for Optimum Properties of Contract DAAJ02-73-C0106. Phase 2, amied at better understanding the critical heat treatment involved in the powder metallurgy production process for manufacturing premium quality hot isostatically pressed As- hip Rene 95 engine hardware. The initial heat treatment study on test blocks and test specimens determined in the effect of section size, quench media, and solution temperature on cooling rate from two different solution temperatures. The correlation between cooling rate microstructure and mechanical property was then established using test specimens. The turbine disks heat treated with different solution temperatures, quench media, and with or without a bore hole defined the most desirable heat treatment parameters for achieving properties. The disks, first heat treated improperly but followed by the selected heat treatment, revealed that the double heat treatment may improve the mechanical properties of the disks initially heat treated improperly. Author (GRA) |