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At least 1,531 records · Page 85
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Turbofan Duct Propagation Model

The CDUCT code utilizes a parabolic approximation to the convected Helmholtz equation in order to efficiently model acoustic propagation in acoustically treated, complex shaped ducts. The parabolic approximation solves one-way wave propagation with a marching method which neglects backwards reflected waves. The derivation of the parabolic approximation is presented. Several code validation cases are given. An acoustic lining design process for an example aft fan duct is discussed. It is noted that the method can efficiently model realistic three-dimension effects, acoustic lining, and flow within the computational capabilities of a typical computer workstation.

Lan, Justin H.↗
NASA NTRS1995

Temperature and Forward Flight Effects on an Underexpanded Jet Noise

An experimental investigation was carried out in the NASA Ames 40'x80' wind tunnel to investigate the far field noise characteristics of a supersonic jet exiting from an axisymmetric convergent nozzle. The nozzle geometry conforms to the ASME specifications for the long radius nozzle. The nozzle pressure ratio NPR, (stagnation pressure/ambient pressure) was varied from 2.5 to 4.5. The temperature ratio, TR (stagnation temperature/ambient temperature) was varied from 2.5 to 3.5. The resulting Reynolds number range, based on the nozzle exit diameter of 12.55cm, is from 0.98 -1.27 x 10(exp 6). The far field directivity was obtained using 1/4 inch diameter condenser microphone with the Flow-Induced Tone Eliminator (FITE) aerodynamic microphone fore body. Appropriate microphone corrections, developed by Allen et al., were made for accurate in-flow acoustic measurements at high frequencies. The narrow band (band width = 64 Hz) frequency spectra covering a range from 0 to 70 KHz were obtained with an accuracy of plus or minus 0.6dB. Typical narrow-band spectra representing far field noise in the forward quadrant, normal to the jet axis and the aft quadrant are shown. The three dominant components of the noise, the screech tone, broad-band shock associated noise and the mixing noise, are clearly identifiable in the spectra. The variation of the screech tone with the exit Mach number (calculated from the isentropic equation relating NPR to Mach number), is shown. The solid line in the figure is calculated using Tam's formula. The model developed by Tam seems to work well even at high temperatures. The appearance of the screech tone is confined to the forward quadrant (theta less than or equal to 60 degrees; the angle theta is measured from the inlet axis). The amplitude of the tone is about 10 dB higher than the broad band noise, which is less than the normally observed amplitude for cold jets operating at similar conditions.

Krothapalli, A.↗
NASA NTRS1996

Features Of Cavity Flow And Acoustics Of Stratospheric Observatory For Infrared Astronomy

The stratospheric observatory for infrared astronomy (SOFIA) is a 2.5 meter aperture Cassegrain telescope with a Nasmyth focus that will be housed in an open cavity in the Boeing 747-SP aircraft and operated at altitudes around 41,000 feet for infrared (IR) viewing of celestial events of astronomical nature. At these altitudes the IR viewing capability of SOFIA far exceeds that of any ground based system. To minimize IR transmission losses, SOFIA will operate with an open cavity. Such an open cavity during flight creates several challenging aerodynamic and aeroacoustic design problems. Foremost of these are: the shear layer over the cavity may cause unwanted resonance if the cavity is untreated; this might give rise to excessive sound pressure levels (SPL) in the cavity and thus affect the unsteady loads on the telescope; the unsteady flow within the cavity produces large dynamic loads and moments that will impact the pointing accuracy of the telescope; the open cavity and the shear layer control devices produce additional drag that will affect directly the time of flight of the mission; the aft location of the cavity down stream of port wing will affect the the flow on the aircraft control surfaces and thus the stability of the aircraft. Also, the highly turbulent shear layer over the cavity and the temperature gradients and 'hot spots' within the cavity can produce a wave front error of the image when it reaches the focal plane of the recorder.

Srinivasan, G. R.↗
NASA NTRS2001

Aerodynamic Performance of Scale-Model Turbofan Outlet Guide Vanes Designed for Low Noise

The design of effective new technologies to reduce aircraft propulsion noise is dependent on an understanding of the noise sources and noise generation mechanisms in the modern turbofan engine. In order to more fully understand the physics of noise in a turbofan engine, a comprehensive aeroacoustic wind tunnel test programs was conducted called the 'Source Diagnostic Test.' The text was cooperative effort between NASA and General Electric Aircraft Engines, as part of the NASA Advanced Subsonic Technology Noise Reduction Program. A 1/5-scale model simulator representing the bypass stage of a current technology high bypass ratio turbofan engine was used in the test. The test article consisted of the bypass fan and outlet guide vanes in a flight-type nacelle. The fan used was a medium pressure ratio design with 22 individual, wide chord blades. Three outlet guide vane design configurations were investigated, representing a 54-vane radial Baseline configuration, a 26-vane radial, wide chord Low Count configuration and a 26-vane, wide chord Low Noise configuration with 30 deg of aft sweep. The test was conducted in the NASA Glenn Research Center 9 by 15-Foot Low Speed Wind Tunnel at velocities simulating the takeoff and approach phases of the aircraft flight envelope. The Source Diagnostic Test had several acoustic and aerodynamic technical objectives: (1) establish the performance of a scale model fan selected to represent the current technology turbofan product; (2) assess the performance of the fan stage with each of the three distinct outlet guide vane designs; (3) determine the effect of the outlet guide vane configuration on the fan baseline performance; and (4) conduct detailed flowfield diagnostic surveys, both acoustic and aerodynamic, to characterize and understand the noise generation mechanisms in a turbofan engine. This paper addresses the fan and stage aerodynamic performance results from the Source Diagnostic Test.

Hughes, Christopher E.↗
NASA NTRS1996

Comparison of Aero/Propulsion Transition Characteristics for a Joint Strike Fighter Configuration

NASA Ames conducted a series tests of a large-scale powered model of Lockheed Martin Skunk Works' JSF STOVL configuration at the 80- by 120-Foot Wind Tunnel. The model uses a modified version of an F-16 fighter engine (a Pratt & Whitney F100-229 engine). To provide vertical flight capability, the Lockheed aircraft incorporates a forward fuselage mounted vertical lift fan in addition to an aft mounted vectoring nozzle all driven by a single engine. The lift fan is driven off the engine by means of a shaft and gear system. The data presented is for STOVL configuration 140, a wing-body-canard planform. The main goal of this testing was to obtain propulsion induced data for evaluating differences between small- and large-scale model testing during the transition flight regime between jet- and wing-borne flight. The data presented is from aerodynamic testing of the model in both the unpowered (engine-off) and powered (engine running) configurations. This data, when combined with the propulsion system thrust calibration data from the NASA Ames' Outdoor Aerodynamic Research Facility (OARF), is used to determine the propulsion induced effects during transition.

Naumowicz, Tim↗
NASA NTRS2002

Integrated Nozzle Design for the GTX RBCC Flowpath

The development of rocket based combined cycle (RBCC) engines are highly dependent upon integrating several different modes of operation into a single system. Due to the integrated nature of the propulsion system, each operating mode relies on the same expansion system to provide thrust. A fixed geometry, altitude-compensating aft-expansion configuration is used for the GTX flowpath configuration. Initial studies on the GTX expansion designs have demonstrated the importance of a smooth, highly integrated design for propulsion system performance. Based upon the results from the initial studies, further design improvements were made to the expansion system. Nozzles designed based on both conical and streamline traced flowfields; are discussed. Results from 3-D CFD calculations on an optimized geometry are also presented. A series of cold-flow experiments are proposed to validate the CFD analysis and quantify performance of the flowpath expansions surface design. A discussion is provided of the research hardware designs and experimental test plans.

Smith, Timothy D.↗
NASA NTRS1986

Coal Research

Coal slurries are "clean" pulverized coal mixed with oil or water. Significant fuel savings can be realized when using coal slurries. Advanced Fuels Technology (AFT) utilized a COSMIC program, (Calculation of Complex Chemical Equilibrium Compositions), which provides specific capabilities for determining combustion products. The company has developed a cleaning process that removes much of the mineral sulphur and ash from the coals.

Source record↗
NASA NTRS2002

Environmentally Compatible Vapor-Phase Corrosion Inhibitor for Space Shuttle Hardware

USA-SRB Element is responsible for the assembly and refurbishment of the non-motor components of the SRB as part of Space Shuttle. Thrust Vector Control (TVC) frames structurally support components of the TVC system located in the aft skirt of the SRB (Solid Rocket Booster). TVC frames are exposed to the seacoast environment after refurbishment and, also, to seawater immersion after splashdown, and during tow-back to CCAFS-Hangar AF refurbishment facilities. During refurbishment operations it was found that numerous TVC frames were experiencing internal corrosion and coating failures, both from salt air and seawater intrusions. Inspectors using borescopes would visually examine the internal cavities of the complicated aluminum alloy welded tubular structure. It was very difficult for inspectors to examine cavity corners and tubing intersections and particularly. to determine the extent of the corrosion and coating anomalies. Physical access to TVC frame internal cavities for corrosion removal and coating repair was virtually impossible, and an improved method using a Liquid (water based) Vapor-phase Corrosion Inhibitor (LVCI) for preventing initiation of new corrosion, and mitigating and/or stopping existing corrosion growth was recommended in lieu of hazardous paint solvents and high VOC/solvent based corrosion inhibitors. In addition, the borescopic inspection method used to detect corrosion, and/or coating anomalies had severe limitations because of part geometry, and an improved non-destructive inspection (NDI) method using Neutron Radiography (N-Ray) was also recommended.

Novak, Howard L.↗
NASA NTRS2002

Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2

This study provides development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the 'open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media 'Darcy' relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations. Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.

Clayton, J. Louie↗
NASA NTRS2002

Prediction of Drag Reduction in Supersonic and Hypersonic Flows with Counterflow Jets

Computational fluid dynamics solutions of the flowfield of a truncated cone-cylinder with and without counterflow jets have been obtained for the short penetration mode (SPM) and long penetration mode (LPM) of the freestream-counterflow jet interaction flowfield. For the case without the counterflow jet, the comparison of the normalized surface pressures showed very good agreement with experimental data. For the case with the SPM jet, the predicted surface pressures did not compare as well with the experimental data upstream of the expansion corner, while aft of the expansion corner, the comparison of the solution and the data is seen to give much better agreement. The difference in the prediction and the data could be due to the transient character of the jet penetration modes, possible effects of the plasma physics that are not accounted for here, or even the less likely effect of flow turbulence, etc. For the LPM jet computations, one-dimensional isentropic relations were used to derived the jet exit conditions in order to obtain the LPM solutions. The solution for the jet exit Mach number of 3 shows a jet penetration several times longer than that of the SPM, and therefore much weaker bow shock, with an attendant reduction in wave drag. The LPM jet is, in essence, seen to be a "pencil" of fluid, with much higher dynamic pressure, embedded in the oncoming supersonic or hypersonic freestream. The methodology for determining the conditions for the LPM jet could enable a practical approach for the design and application of counterflow LPM jets for the reduction of wave drag and heat flux, thus significantly enhancing the aerodynamic characteristics and aerothermal performance of supersonic and hypersonic vehicles. The solutions show that the qualitative flow structure is very well captured. The obtained results, therefore, suggest that counterflowing jets are viable candidate technology concepts that can be employed to give significant reductions in wave drag, heat flux, and other attendant aerodynamic benefits.

Daso, Endwell O.↗
NASA NTRS2003

Fan Noise Control Using Herschel-quincke Resonators

The research effort proposed for this NASA NRA is mainly experimental. In addition, Virginia Tech is working in partnership with Goodrich Aerospace, Aerostructures Group for the analytical development needed to support the experimental endeavor, i.e. model development, design, and system studies. In this project, Herschel-Quincke (HQ)liner technology experiments will be performed at the NASA Glenn Active Noise Control Fan (ANCF) facility. A schematic of both inlet and aft HQ-liner systems installed in the ANCF rig as well as a picture of the Glenn facility is shown. The main goal is to simultaneously test in both the inlet and bypass duct sections. The by-pass duct will have HQ-systems in both the inner and outer duct walls. The main advantages of performing tests at the ANCF facility are that the effect of the inlet HQ-system on the by-pass HQ-system and vice versa, can be accurately determined from the in-duct modal data. Another significant advantage is that it offers the opportunity to assess (on a common basis) the proposed noise reduction concept on the ANCF rig which in the past has been used for assessing other active and passive noise reduction strategies.

Burdisso, Ricardo A.↗
NASA NTRS1947

Stick-Fixed Stability and Control Characteristics of the Consolidated Vultee Model 240 Airplane as Estimated from Tests of a 0.092-Scale Powered Model

Estimates of the static stick-fixed stability and control characteristics of the Consolidated Vultee model 240 airplane are presented in this report. The estimates are based on tests of a 0.092-scale powered model in the 10-foot wind tunnel of the Guggenheim Aeronautical Laboratory of the California Institute of Technology. Results of the analysis are evaluated in terms of the Army specifications for stability and control characteristics which are more specific and, in general, equal to or more rigid than the Civil Aeronautics Administration requirements. The stick-fixed stability and control characteristics of the Consolidated Vultee model 240 were found to be satisfactory except for the following: 1) Marginal longitudinal stability in the landing approach (flaps 30 deg, 50% minimum continuous power) with aft center of gravity (31% M.A.C.); 2) Marginal rudder control to hold zero sideslip in a climb after take-off with asymmetric power (flaps 30 deg, left engine inoperative, right engine delivering take-off power) with maximum rudder throw limited to +/- 18 deg; 3) Marginal dihedral effect with flaps 40 deg and engines delivering maximum continuous power.

McCullough, George B.↗
NASA NTRS2003

The Development of the Ducted Fan Noise Propagation and Radiation Code CDUCT-LaRC

The development of the ducted fan noise propagation and radiation code CDUCT-LaRC at NASA Langley Research Center is described. This code calculates the propagation and radiation of given acoustic modes ahead of the fan face or aft of the exhaust guide vanes in the inlet or exhaust ducts, respectively. This paper gives a description of the modules comprising CDUCT-LaRC. The grid generation module provides automatic creation of numerical grids for complex (non-axisymmetric) geometries that include single or multiple pylons. Files for performing automatic inviscid mean flow calculations are also generated within this module. The duct propagation is based on the parabolic approximation theory of R. P. Dougherty. This theory allows the handling of complex internal geometries and the ability to study the effect of non-uniform (i.e. circumferentially and axially segmented) liners. Finally, the duct radiation module is based on the Ffowcs Williams-Hawkings (FW-H) equation with a penetrable data surface. Refraction of sound through the shear layer between the external flow and bypass duct flow is included. Results for benchmark annular ducts, as well as other geometries with pylons, are presented and compared with available analytical data.

Nark, Douglas M.↗
NASA NTRS2003

RSRM and ETM03 Internal Flow Simulations and Comparisons

ETM03 (Engineering Test Motor-03) is an extended length RSRM (Reusable Solid Rocket Motor) designed to increase motor performance and create more severe internal environments compared with the standard four-segment RSRM motor configuration. This is achieved primarily through three unique design features. First is the incorporation of an additional RSRM center segment, second is a slight increase in throat diameter, and third is the use of an Extended Aft Exit Cone (EAEC). As a result of these design features, parameters such as web time, action time, head end pressure, web time average pressure, maximum thrust, mass flow rate, centerline Mach number, pressure and thrust integrals have all increased compared with nominal RSRM values. In some cases these increases are substantial. The primary objective of the ETM03 test program is to provide a platform for RSRM component margin testing. Test results will not only provide direct data concerning component performance under more adverse conditions, but serve as a second design data point for developing, validating and enhancing component analytical modeling techniques. To help component designers assess how the changes in motor environment will affect performance, internal flow simulations for both the nominal RSRM and ETM03 motor designs were completed to obtain comparisons of aero-thermal boundary conditions and system loads. Full geometries for both motors were characterized with two-dimensional axi-symmetric models at burn times of 1, 20, 54, 67 and 80-seconds. A sixth set considered burn times of 110 and 117-seconds for RSRM and ETM03, respectively. The simulations were performed using the computational fluid dynamics (CFD) commercial code FLUENT (trademark). Of particular interest were any differences between the two motor environments that could lead to a significant increase in system loads, or in internal insulation and/or nozzle component charring and erosion in ETM03 compared with RSRM. Based on these comparative analyses conducted in this study, the objective of ETM03 will be achieved by providing a more adverse operating environment for motor components than the nominal RSRM environment. For example: Higher chamber pressure drop in ETM03 than in RSRM; higher centerline Mach numbers approaching the nozzle in ETM03 than in RSRM; higher heat transfer rates for the internal insulation and nozzle components in ETM03 than in RSRM; and higher levels of droplet impingement and slag accumulation in ETM03 than in the RSRM.

Ahmad, R. A.↗
NASA NTRS2003

Comparison of Acoustic Impedance Eduction Techniques for Locally-Reacting Liners

Typical acoustic liners used in current aircraft inlets and aft-fan ducts consist of some type of perforated facesheet bonded to a honeycomb core. A number of techniques for determining the acoustic impedance of these locallyreacting liners have been developed over the last five decades. In addition, a number of models have been developed to predict the acoustic impedance of locallyreacting liners in the presence of grazing flow, and to use that information together with aeroacoustic propagation codes to assess the noise absorption provided by these liners. These prediction models have incorporated the results from databases acquired with specific impedance eduction techniques. Thus, while these prediction models are acceptable for liners that are similar to those tested in these databases, their application to new liner configurations must be viewed with caution. The primary purpose of this paper is to provide a comparison of impedance eduction techniques that have been implemented at various aerospace research laboratories in the United States (NASA Langley Research Center, General Electric Aircraft Engines, B. F. Goodrich and Boeing). A secondary purpose is to provide data for liner configurations that extend the porosity range beyond that which has been previously used in common aircraft engine nacelles. Two sets of liners were designed to study the effects of three parameters: perforate hole diameter, facesheet thickness and porosity. These two sets of liners were constructed for testing in each of the laboratories listed above. The first set of liners was designed to fit into the NASA Langley and Boeing test facilities. The second set was designed to fit into the General Electric Aircraft Engines and B. F. Goodrich test facilities. By using the same parent material, both sets of liners were identical to within the limits of material and fabrication variability. Baseline data were obtained in the normal incidence impedance tubes at NASA Langley and B. F. Goodrich. The results were found to compare extremely well. The samples were then tested in the grazing flow ducts of each of the four laboratories. Perhaps the most significant result of these comparisons is that the educed acoustic resistances for the liners used in this study increase as the mean flow profile is modified from uniform to 3-D shear. This realization has demonstrated the need for an frequency-dependent impedance eduction technique that incorporates 3-D shear flow and is efficient.

Jones, M. G.↗
NASA NTRS1956

On Landing Gear Stresses

Information on landing gear stresses is presented on the following: vibratory phenomena, tangential forces applied to landing gear, fore and aft oscillations of landing gears, examples of fatigue failures, vibration calculations, and improvement of existing test equipment.

Gentric, A.↗
NASA NTRS2003

Development of Displacement Gages Exposed to Solid Rocket Motor Internal Environments

The Space Shuttle Reusable Solid Rocket Motor (RSRM) has three non-vented segment-to-segment case field joints. These joints use an interference fit J-joint that is bonded at assembly with a Pressure Sensitive Adhesive (PSA) inboard of redundant O-ring seals. Full-scale motor and sub-scale test article experience has shown that the ability to preclude gas leakage past the J-joint is a function of PSA type, joint moisture from pre-assembly humidity exposure, and the magnitude of joint displacement during motor operation. To more accurately determine the axial displacements at the J-joints, two thermally durable displacement gages (one mechanical and one electrical) were designed and developed. The mechanical displacement gage concept was generated first as a non-electrical, self-contained gage to capture the maximum magnitude of the J-joint motion. When it became feasible, the electrical displacement gage concept was generated second as a real-time linear displacement gage. Both of these gages were refined in development testing that included hot internal solid rocket motor environments and simulated vibration environments. As a result of this gage development effort, joint motions have been measured in static fired RSRM J-joints where intentional venting was produced (Flight Support Motor #8, FSM-8) and nominal non-vented behavior occurred (FSM-9 and FSM-10). This data gives new insight into the nominal characteristics of the three case J-joint positions (forward, center and aft) and characteristics of some case J-joints that became vented during motor operation. The data supports previous structural model predictions. These gages will also be useful in evaluating J-joint motion differences in a five-segment Space Shuttle solid rocket motor.

Bolton, D. E.↗
NASA NTRS2003

TES Limb-Geometry Observations of Aerosols

The Thermal Emission Spectrometer (TES) on-board Mars Global Surveyor (MGS) has a pointing mirror that allows observations in the plane of the orbit anywhere from directly nadir to far above either the forward or aft limbs for details about the TES instrument). Nadir-geometry observations are defined as those where the field-of-view contains the surface of Mars (even if the actual observation is at a high emission angle far from true nadir). Limb-geometry observations are defined as those where the line-of-sight of the observations does not intersect the surface. At a number of points along the MGS orbit (typically every 10 deg. or 20 deg. of latitude) a limb sequence is taken, which includes a stack of overlapping TES spectra from just below the limb to more than 120 km above the limb. A typical limb sequence has approx. 20 individual spectra, and the projected size of a TES pixel at the limb is 13 km.

Michael D Smith↗