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Detailed information regarding the design and installation of a heating and hot water system in a commercial application is given. This information includes descriptions of system and building, design philosophy, control logic operation modes, design and installation drawing and a brief description of problems encountered and their solutions.
Relevant differences between the MPPM resident IBM 370computer and the NASA Sigma 9 computer are described as well as the MPPM system itself and its development. Problems encountered and solutions used to overcome these difficulties during installation of the MPPM system at MSFC are discussed. Remaining work on the installation effort is summarized. The relevant hardware features incorporated in the program are described and their implications on the transportability of the MPPM source code are examined.
Investigations have been made in flight and in wind tunnels to determine which components of turbojet installations are most critical in icing conditions, and to evaluate several methods of icing protection. From these studies, the requirements necessary for adequate icing protection and the consequent penalties on engine performance can be estimated. Because investigations have indicated that the compressor-inlet screen constitutes the greatest icing hazard and is difficult to protect, complete removal or retraction of the screen upon encountering an icing condition is recommended. In the absence of the screen, the inlet guide vanes of an axial-flow-type turbojet engine constitute the greatest danger to engine operation in an icing condition; a centrifugal-type engine, on the other hand, is relatively unsusceptible to icing once the screen has been removed. Of the three icing-protection systems investigated, surface heating, hot-gas bleedback, and inertia-separation inlets, only the first two offer an acceptable solution to the problem of engine icing protection. Surface heating, either by gas heating or electrical means, appears to be the most acceptable icing-protection method with regard to performance losses. Hot-gas bleedback, although causing undesirable thrust losses, offers an easy means of obtaining icing protection for some installations. The final choice of an icing-protection system depends, however, on the supply of heated gas and electrical power available and on the allowable performance and. weight penalties associated with each system.
The design of the propulsion system installation affects strongly the total drag and overall performance of an aircraft, and the concept, placement, and integration details of the exhaust nozzle are major considerations in the configuration definition. As part of the NASA Supersonic Cruise Research (SCR) program, a wind tunnel test program has been conducted to investigate exhaust nozzle-airframe interactions at transonic speeds. First phase testing is to establish guidelines for follow-on testing. A summary is provided of the results of first phase testing, taking into account the test approach, the effect of nozzle closure on aircraft aerodynamic characteristics, nozzle installation effects and nacelle interference drag, and an analytical study of the effects of nozzle closure on the aircraft.
The aeropropulsive characteristics of an advanced twin engine fighter designed for supersonic cruise was investigated in the 16 foot Transonic Tunnel. The performance characteristics of advanced nonaxisymmetric nozzles installed in various nacelle locations, the effects of thrust induced forces on overall aircraft aerodynamics, the trim characteristics, and the thrust reverser performance were evaluated. The major model variables included nozzle power setting; nozzle duct aspect ratio; forward, mid, and aft nacelle axial locations; inboard and outboard underwing nacelle locations; and underwing and overwing nacelle locations. Thrust vectoring exhaust nozzle configurations included a wedge nozzle, a two dimensional convergent divergent nozzle, and a single expansion ramp nozzle, each with deflection angles up to 30 deg. In addition to the nonaxisymmetric nozzles, an axisymmetric nozzle installation was also tested. The use of a canard for trim was also assessed.
A new electronic scanner of pressure (ESOP) has been developed by NASA Ames Research Center for installation in wind-tunnel models. An ESOP system includes up to 20 pressure modules, each with 48 pressure transducers, an A/D converter, a microprocessor, a data controller, a monitor unit, and a heater controller. The system is sized so that the pressure modules and A/D converter module can be installed within an average-size model tested in the Ames Aerodynamics Division wind tunnels. This paper describes the ESOP system, emphasizing the main element of the system - the pressure module. The measured performance of the overall system is also presented.
A trend toward replacement of parametric model testing with parametric analysis for the design of aircraft is driven by the rapidly escalating cost of wind tunnel testing, the increasing availability of large fast computers, and powerful numerical flow algorithms. In connection with the complex flow phenomena characteristic of propulsion installations, it is now necessary to employ both parametric analysis and testing for design procedures. Powerful flow analysis techniques are available to predict local flow phenomena. However, the employment of these techniques is very expensive. It is, therefore, necessary to link these analyses with less powerful and less expensive procedures for an accurate analysis of propulsion installation flowfields. However, the interfacing and coupling processes needed are not available. The present investigation is concerned with progress made regarding the development of suitable linking methods. Attention is given to methods of analysis for predicting the flow around a nacelle coupled to a highly swept wing.
Methods and devices for installing in a tubular shaft multiple strain gages are disclosed with focus on a method and a device for pneumatically forcing strain gages into seated engagement with the internal surfaces of a tubular shaft in an installation of multiple strain gages in a tubular shaft. The strain gages or other electron devices are seated in a template-like component which is wrapped about a pneumatically expansible body. The component is inserted into a shaft and the body is pneumatically expanded after a suitable adhesive was applied to the surfaces.
The NASA Lewis Research Center in cooperation with the World Health Organization, U.S.A. I.D., the Pan American Health Organization and national government agencies in some developing countries sponsored the installation of twenty photovoltaic powered medical vaccine storage refrigerator-freezer (R/F) systems. The Solar Power Corporation was selected as the contractor to perform the design, development and installation of these twenty units. Solar Power's experiences are described herein.
The installation procedure, maintenance, adjustment and operation of a Lucas type fuel injection system for 13B rotary racing engine is outlined. Components of the fuel injection system and installation procedure and notes are described. Maintenance, adjustment, and operation are discussed.
In order to understand the effects of installation on propeller noise, numerous measurements are required to define the directivity of the noise as well as the level. An experimental study was designed to map the noise radiation pattern for various single-rotation propeller (SRP) and counter-rotation propeller (CRP) installations covering + or 60 deg from the propeller disk plane and + or - 60 deg laterally. Configurations that were considered included an SRP at angle of attack and in tractor and pusher operations and a CRP. A first principles linear theory was validated for the SRP tractor operation over the angle range mentioned above. The increases in noise that arise from an unsteady loading operation such as an SRP pusher or CRP exceed 15 dB and depend on the observer location. In particular, the majority of the additional noise appears to radiate in the axial directions.
An experimental investigation was recently conducted in the NASA-Langley Research Center 4- by 7-Meter Tunnel to explore the effects of aft-mounted advanced turboprop installations on the low-speed stability and control characteristics of a representative transport aircraft configuration. For the conditions investigated, the experimental results indicate that the longitudinal and directional stability of the configuration was not adversely affected by aft-pylon-mounted single-rotation tractor or counter-rotation pusher propeller installations. The results indicate that for the single-rotation tractor configuration the propeller induced aerodynamics are dominated by the interaction of the propeller slipsteam with the pylon, and that the propeller contribution to the stability characteristics for the counter-rotating pusher configuration is dominated by propeller normal forces. The longitudinal and directional control effectiveness, engine-out characteristics, and ground effects are also presented.
The installation interference effects of an underwing-mounted, long duct, turbofan nacelle were evaluated in the Langley 16-Foot Transonic Tunnel with two different pylon shapes installed on a twin engine transport model having a supercritical wing swept 30 deg. Wing, pylon, and nacelle pressures and overall model force data were obtained at Mach numbers from 0.70 to 0.83 and nominal angles of attack from -2 deg to 4 deg at an average unit Reynolds number of 11.9 x 1,000,000 per meter. The results show that adding the long duct nacelles to the supercritical wing, in the near sonic flow field, changed the magnitude and direction of flow velocities over the entire span, significantly reduced cruise lift, and caused large interference drag on the nacelle afterbody.
Measurements which are required to define the directivity and the level of propeller noise were studied. The noise radiation pattern for various single-rotation (SR) propeller and counter-rotation (CR) propeller installations were mapped. The measurements covered + or - 60 deg from the propeller disk plane and + or - 60 deg in the cross-stream direction. Configurations examined included SR and CR propellers at angle of attack and an SR pusher installation. The increases in noise that arise from an unsteady loading operation such as an SR pusher or a CR exceeded 15 dB in the forward axial direction. Most of the additional noise radiates in the axial directions for unsteady loading operations of both the SR pusher and the CR tractor.
A simple barrier shielding model was used to estimate the amount of noise shielding on the fuselage that could result from installing a short duct around a wing-mounted advanced propeller. With the propeller located one-third of the duct length from the inlet, estimates for the maximum blade passing tone attenuation varied from 7 dB for a duct 0.25 propeller diameter long to 16.75 dB for a duct 1 diameter long. Attenuations for the higher harmonics would be even larger because of their shorter wavelengths relative to the duct length. These estimates show that the fuselage noise reduction potential of a ducted compared with an unducted propeller is significant. Even more reduction might occur if acoustic attenuation material were installed in the duct.
The three-dimensional transonic Euler method with boundary layer interaction is used to study the flow about a transport aircraft equipped with NASA-developed compression pylons. The results show that the present pylons perform well under the installed conditions, reducing the wing/pylon junction velocities and thus reducing the installed loft loss and drag. The predicted theoretical results are found to agree moderately well with experimental wind tunnel results.
A laboratory-based test apparatus was employed to investigate the effects of power-plant placement, engine/nacelle mass installation, and wing-to-fuselage attachment methods on propeller-induced structure-borne noise (SBN) transmission levels and their effects on noise-control measures. Data are presented showing SBN transmission is insensitive to propeller spanwise placement, however some sensitivity is seen in propeller-to-wing spacing. Installation of an engine/nacelle mass and variation in wing-to-fuselage attachments have measurable influences on SBN transmission and control measures.
The test section of the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel was acoustically treated to allow the measurement of sound under simulated free-field conditions. The treatment was designed for high sound absorption at frequencies above 250 Hz and for withstanding the environmental conditions in the test section. In order to achieve the design requirements, a fibrous, bulk-absorber material was packed into removable panel sections. Each section was divided into two equal-depth layers packed with material to different bulk densities. The lower density was next to the facing of the treatment. The facing consisted of a perforated plate and screening material layered together. Sample tests for normal-incidence acoustic absorption were also conducted in an impedance tube to provide data to aid in the treatment design. Tests with no airflow, involving the measurement of the absorptive properties of the treatment installed in the 9- by 15-foot wind tunnel test section, combined the use of time-delay spectrometry with a previously established free-field measurement method. This new application of time-delay spectrometry enabled these free-field measurements to be made in nonanechoic conditions. The results showed that the installed acoustic treatment had absorption coefficients greater than 0.95 over the frequency range 250 Hz to 4 kHz. The measurements in the wind tunnel were in good agreement with both the analytical prediction and the impedance tube test data.