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NASA NTRS1975

Pressure distribution in a converging-diverging nozzle during two-phase choked flow of subcooled nitrogen

Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane is at the divergence end of the constant area throat section. At high stagnation pressures the fluid stays liquid well into the constant area throat region; at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region, and this anomaly is related to the two-phase process. The fluid is metastably all liquid below the saturation pressure.

Simoneau, R. J.↗

NASA NTRS1975

Pressure distribution in a converging-diverging nozzle during two-phase choked flow of subcooled nitrogen

Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane appears to be at the divergence end of the constant area throat section. At high stagnation pressures the fluid appears to stay liquid well into the constant area throat region; however, at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region. This anomaly appears to be fundamentally related to the two-phase process and not merely to the present specific nozzle geometry. The fluid appears to be metastably all liquid below the saturation pressure. The data are compared to various flow models. No model adequately describes the whole range of the experiment.

Simoneau, R. J.↗

NASA NTRS2009

Fault Diagnostics and Prognostics for Large Segmented SRMs

We report progress in development of the fault diagnostic and prognostic (FD&P) system for large segmented solid rocket motors (SRMs). The model includes the following main components: (i) 1D dynamical model of internal ballistics of SRMs; (ii) surface regression model for the propellant taking into account erosive burning; (iii) model of the propellant geometry; (iv) model of the nozzle ablation; (v) model of a hole burning through in the SRM steel case. The model is verified by comparison of the spatially resolved time traces of the flow parameters obtained in simulations with the results of the simulations obtained using high-fidelity 2D FLUENT model (developed by the third party). To develop FD&P system of a case breach fault for a large segmented rocket we notice [1] that the stationary zero-dimensional approximation for the nozzle stagnation pressure is surprisingly accurate even when stagnation pressure varies significantly in time during burning tail-off. This was also found to be true for the case breach fault [2]. These results allow us to use the FD&P developed in our earlier research [3]-[6] by substituting head stagnation pressure with nozzle stagnation pressure. The axial corrections to the value of the side thrust due to the mass addition are taken into account by solving a system of ODEs in spatial dimension.

Luchinsky, Dmitry↗

NASA NTRS1996

Investigation of Condensation/Clustering Effects on Rayleigh Scattering Measurements in a Hypersonic Wind Tunnel

Rayleigh scattering, a nonintrusive measurement technique for the measurement of density in a hypersonic wind tunnel, is under investigation at Wright Laboratory's Mach 6 wind tunnel. Several adverse effects, i.e., extraneous scatter off walls and windows, hinder Rayleigh scattering measurements. Condensation and clustering of flow constituents also present formidable obstacles. Overcoming some of these difficulties, measurements have been achieved while the Mach 6 test section was pumped down to a vacuum, as well as for actual tunnel operation for various stagnation pressures at fixed stagnation temperatures. Stagnation pressures ranged from 0.69 MPa to 6.9 MPa at fixed stagnation temperatures of 511, 556, and 611 K. Rayleigh scatter results show signal levels much higher than expected for molecular scattering in the wind tunnel. Even with higher than expected signals, scattering measurements have been made in the flowfield of an 8-degree half-angle blunt nose cone with a nose radius of 1.5 cm.

Tyler, Charles↗

NASA NTRS1994

The NASA Langley 8-foot Transonic Pressure Tunnel calibration

The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.

Brooks, Cuyler W., Jr.↗

NASA NTRS1987

Digital program for calculating static pressure position error

A computer program written to calculate the static pressure position error of airspeed systems contains five separate methods for determining position error, of which the user may select from one to five at a time. The program uses data from both the test aircraft and the ground-based radar to calculate the error. In addition, some of the methods require rawinsonde data or an atmospheric analysis, or both. The program output lists the corrections to Mach number, altitude, and static pressure that are due to position error. Reference values such as angle of attack, angle of sideslip, indicated Mach number, indicated pressure altitude, stagnation pressure, and total temperature are also listed.

Johnson, J. Blair↗

NASA NTRS1989

The NASA Langley Research Center 0.3-meter transonic cryogenic tunnel T-P/Re-M controller manual

A new microcomputer based controller for the 0.3-m Transonic Cryogenic Tunnel (TCT) has been commissioned in 1988 and has reliably operated for more than a year. The tunnel stagnation pressure, gas stagnation temperature, tunnel wall structural temperature and flow Mach number are precisely controlled by the new controller in a stable manner. The tunnel control hardware, software, and the flow chart to assist in calibration of the sensors, actuators, and the controller real time functions are described. The software installation details are also presented. The report serves as the maintenance and trouble shooting manual for the 0.3-m TCT controller.

Balakrishna, S.↗

NASA NTRS1949

Performance of J33 turbojet engine with shaft-power extraction III : turbine performance

The performance of the turbine component of a J33 turbojet engine was determined over a range of turbine speeds from 8000 to 11,500 rpm.Turbine-inlet temperature was varied from the minimum required to drive the compressor to a maximum of approximately 2000 degrees R at each of several intermediate turbine speeds. Data are presented that show the horsepower developed by the turbine per pound of gas flow. The relation between turbine-inlet stagnation pressure, turbine-outlet stagnation pressure, and turbine-outlet static pressure was established. The turbine-weight-flow parameter varied from 39.2 to 43.6. The maximum turbine efficiency measured was 0.86 at a pressure ratio of 3.5 and a ratio of blade speed to theoretical nozzle velocity of 0.39. A generalized performance map of the turbine-horsepower parameter plotted against the turbine-speed parameter indicated that the best turbine efficiency is obtained when the turbine power is 10 percent greater than the compressor horsepower. The variation of efficiency with the ratio of blade speed to nozzle velocity indicated that the turbine operates at a speed above that for maximum efficiency when the engine is operated normally with the 19-inch-diameter jet nozzle.

Huppert, M C↗

NASA NTRS1992

Design and evaluation of single and dual flow thrust vector nozzles with post exit vanes

This Thrust Vectored Research project required that a 1/24 scale model of the F/A-18 High Alpha Research Vehicle, (HARV), propulsion system be constructed on the university campus. This propulsion system was designed for cold flow testing on a multicomponent test rig. Forces and moments were measured to study nozzle performance parameters. The flow visualization technique of color Schlieren photography was performed to investigate the flow phenomena at the nozzle exit. The flow interactions that were identified consisted of vane nozzleing between the outer and lower vanes and vane tip interference. The thrust vectoring system consisted of three asymmetrically spaced vanes installed circumferentially on a maximum afterburner nozzle. The performance of the nozzle was investigated with the outer and lower vanes equally deflected, (-10 deg is less than delta(sub v) is less than 25 deg), and with the upper vane fully retracted, (delta(sub v) equals -10 deg). The nozzle pressure ratio ranged from 4 to 6. The results indicated that a vane nozzleing effect developed at nozzle pressure ratios of 4 and 6 when the outer and lower vanes were deflected far enough into the flow field such that the increase in vane area accelerated the flow past the vanes causing distorted shock waves. This accelerated flow was a result of a pressure differential existing between the inside surface of the vane and the ambient pressure. The stagnation pressure that developed along the inside surface of the vane accelerated the flow past the vanes causing it to equalize with ambient pressure, thus providing distorted shock waves. A tip interference was present at the trailing edge of the upper vane as a result of low nozzle pressure, NPR 4, with high vane deflection, delta(sub v) equals 25 degrees, and also with a high nozzle pressure, NPR 6, and low vane deflections, delta(sub v) equals 15 degrees.

Carpenter, Thomas W.↗

NASA NTRS1986

Computation of nozzle flow fields using the PARC2D Navier-Stokes code

Supersonic nozzles which operate at low Reynolds numbers and have large expansion ratios have very thick boundary layers at their exit. This leads to a very strong viscous/inviscid interaction upon the flow within the nozzle and the traditional nozzle design techniques which correct the inviscid core with a boundary layer displacement do not accurately predict the nozzle exit conditions. A full Navier-Stokes code (PARC2D) was used to compute the nozzle flow field. Grids were generated using the interactive grid generator code TBGG. All computations were made on the NASA MSFC CRAY X-MP computer. Comparison was made between the computations and in-house wall pressure measurements for CO2 flow through a conical nozzle having an area ratio of 40. Satisfactory agreement existed between the computations and measurements for a stagnation pressure of 29.4 psia and stagnation temperature of 1060 R. However, agreement did not exist at a stagnation pressure of 7.4 psia. Several reasons for the lack of agreement are possible. The computational code assumed a constant gas gamma whereas gamma for CO2 varied from 1.22 in the plenum chamber to 1.38 at the nozzle exit. Finally, it is possible that condensation occurred during the expansion at the lower stagnation pressure.

Collins, Frank G.↗

NASA NTRS1995

Ames Research Center Mars/Pathfinder Heat Shield Design Verification ARC-JET Test

Design verification tests were performed on samples representing the aerobrake of the Mars/Pathfinder vehicle. The test specimens consisted of the SLA-561V ablator bonded to the honeycomb structure. The primary objective was to evaluate the ablation materials performance and to measure temperatures within the ablator, at the structural bondline and at the back sheet of the honeycomb structure. Other objectives were to evaluate the effect of ablative repair plug material treatment and voids in the heat shield. A total of 29 models were provided for testing in the Ames 60MW arc-jet facility. Of these, 23 models were flat-faced and six remaining models were curved edge ones, intended to simulate the conditions on the curved rim of the forebody where the maximum shear occurred. Eight sets of test conditions were used. The stagnation point heating rates varied from 47 to 240 W/cm2 and the stagnation pressures from 0.15 to 0.27 atm. (The maximum flight values are 132 W/cm2 and 0.25 atm) The majority of these runs were made at a nominal stagnation pressure of 0.25 atm. Two higher pressure runs were made to check the current (denser) ablation material for spallation, or other forms of thermal stress failure. Over 60% of the flatfaced models yielded good thermocouple data and all produced useful surface recession information. Of the five curved-edge models that were tested, only one gave good data; the remaining ones experienced model-holder failure. The test results can be summarized by noting that no failure of the ablative material was observed on any model. Also, the bondline temperature design limit of 250 C was never reached within an equivalent flight time despite a stagnation point heat load that exceeded the maximum flight value by up to 130%. At heating rates of over 200W/cm2 and stagnation pressures of 0.25 atm, or greater, the average surface recessions exceeded 0.5 cm on some models. The surface roughness increased dramatically at pressures above 0.25 atm and was four times greater at 0.27 atm than at 0.25 atm. Procured repair plug material performed much better than room-temperature cured plugs, as observed in the previous tests. Voids in the ablator did not increase local temperatures and gaps did not grow during testing.

Tran, Huy K.↗

NASA NTRS1980

Inlet flow distortion in turbomachinery. I - Comparison of theory and experiment in a transonic fan stage. II - A parameter study

In the present paper, a semi-actuator-disk theory is reviewed that was developed previously for the distorted inflow to a single-stage axial-flow compressor. Flow distortion occurs far upstream; it may be a distortion in stagnation temperature, stagnation pressure, or both. Losses, quasi-steady deviation angles, and reference incidence correlations are included in the analysis, and both subsonic and transonic relative Mach numbers are considered. The theory is compared with measurements made in a transonic fan stage, and a parameter study is carried out to determine the influence of solidity on the attenuation of distortions in stagnation pressure and stagnation temperature.

Seidel, B. S.↗

NASA NTRS1974

Operating characteristics of the Langley Mach 10 high Reynolds number helium tunnel

Operating characteristics of the Langley Mach 10 high Reynolds number helium tunnel are presented for stagnation pressures from 138 N/sq cm to 1655 N/sq cm. The characteristics include detailed Mach number surveys in the test section from which usable core size and regions of disturbed flow were determined, preliminary blockage test results, and maximum run time to be expected at various stagnation pressures. Important tunnel dimensions including details of the model mounting apparatus are given. Measurements show the variation in average core Mach number in the test section to be between 9.4 and 10 for the present range of test conditions. The core radius is from 23 cm to 31.5 cm, depending on stagnation pressure and axial location in the test section.

Watson, R. D.↗

NASA NTRS1974

Flight simulation characteristics of the Langley high Reynolds number cryogenic transonic tunnel

The characteristics of the Langley 34 cm (13.5 in.) pilot cryogenic transonic pressure tunnel are described, and the results of initial tunnel operation are presented. Tests of a two-dimensional airfoil at a Mach number of 0.85 show identical pressure distributions for a chord Reynolds number of 8,600,000 obtained first at a stagnation pressure of 4.91 atmospheres at a stagnation temperature of +120 F and then at a stagnation pressure of 1.19 atmospheres at a stagnation temperature of -250 F.

Kilgore, R. A.↗

NASA NTRS1974

Simulation of flight test conditions in the Langley pilot transonic cryogenic tunnel

The theory and advantages of the cryogenic tunnel concept are briefly reviewed. The unique ability to vary temperature independently of pressure and Mach number allows, in addition to large reductions in model loads and tunnel power, the independent determination of Reynolds number, Mach number, and aeroelastic effects on the aerodynamic characteristics of the model. Various combinations of Reynolds number and dynamic pressure are established to represent accurately flight variations of aeroelastic deformation with altitude changes. The consequences of the thermal and caloric imperfections of the test gas under cryogenic conditions were examined and found to be insignificant for operating pressures up to 5 atm. The characteristics of the Langley pilot transonic cryogenic tunnel are described and the results of initial tunnel operation are presented. Tests of a two-dimensional airfoil at a Mach number of 0.85 show identical pressure distributions for a chord Reynolds number of 8,600,000 obtained first at a stagnation pressure of 4.91 atm at a stagnation temperature of 322.0 K and then at a stagnation pressure of 1.19 atm at a stagnation temperature of 116.5 K.

Kilgore, R. A.↗

NASA NTRS1974

An exploratory investigation of a wake disruption technique for studying wake reestablishment time

An exploratory investigation was made of a wake disruption technique for studying the hypersonic-wake reestablishment time in a blowdown wind tunnel. In this technique, a highly underexpanded jet issuing from the base of a 10 deg half-angle cone totally disrupts and displaces the conventional wake. The jet was rapidly shut off by an explosively actuated valve and the time for wake reestablishment was measured. The tests were conducted in the Mach 6 high Reynolds number tunnel at a stagnation temperature of 506 K and stagnation pressure of 2.86 MPa. The model base jet stagnation pressure was 3.55 MPa at room temperature. High-speed schlieren motion pictures indicated that disappearance of the disrupting jet and reestablishment of the wake-recompression shock were probably occurring simultaneously and that the time disruptive-jet-air shutoff to wake recompression shock reestablishment was probably between 200 and 450 microseconds (flow lengths from 1.8 to 4.2). The values of flow lengths are about one-thord to one-half the values measured in impulse facilities in a previous study. This shorter time is believed to be largely due to difference in flow conditions between the jet disruption technique and impulse facilities.

Clark, L. E.↗

NASA NTRS1996

PICA Forebody Heatshield Qualification for the Stardust Discovery Class Mission

This paper presents the qualification of the light weight Phenolic Impregnated Carbon Ablators (PICA) as the forebody heatshield for the Stardust Discovery Class Mission. The Stardust spacecraft will be launched in early 1999 and fly by Comet Wild-2 to collect cometary and interstellar dust and return them back to earth in the Sample Return Capsule (SRC). This earth re-entry will be the fastest to date, at 12.6 km/s, and therefore requires a heatshield that can withstand very high heating rates and stagnation pressures, as well as provide the necessary insulation to the vehicle structure. The PICA material was developed as part of the Lightweight Ceramic Ablators program at NASA Ames Research Center, and was baselined as the forebody heatshield because of its low density and superior ablation and thermal performance at severe aerothermodynamic conditions. Under a Small Business Innovative Research (SBIR) program with NASA Ames, Fiber Materials, Inc. developed a process to manufacture a single-piece PICA heatshield for the forebody of the SRC, along with witness material for the fabrication of the test models. The test models were fabricated and instrumented by the staff of Lockheed Martin Astronautics in Denver, Colorado. Full body preliminary aerothermal CFD calculations were performed at NASA Ames to determine the heating and stagnation pressure conditions. The Heat shield sizing was also performed at NASA Ames by using a new material response code that accounts for the highly porous characteristics of the PICA material. The ablation and thermal performance of PICA was qualified in the NASA Ames Interaction Heating Arc Jet Facility. A total of 24 models and four test conditions were used to qualify PICA at the predicted peak heat flux, heat load, shear, and stagnation pressure conditions. Surface and in-depth temperatures were measured using optical pyrometers and thermocouples. Surface recession was measured by using a template and a height gage. Several models were tested to evaluate repair procedures, and two models were cold soaked in liquid nitrogen, prior th testing, to investigate the effect of the cold space environment on the performance of the material. In addition, material cored from a demonstration single-piece heatshield was tested to verify that the PICA process can be successfully completed on a large, complex heatshield shape.

Tran, Huy K.↗

NASA NTRS1993

Flow field measurements in a crossing shockwave turbulent boundary layer interaction at Mach 3

An experimental study has been conducted to examine the flow field of the 3D crossing shock wave/turbulent boundary layer interaction. A symmetric pair of 9-deg fins were used to generate the crossing shocks. The incoming boundary layer was developed on the tunnel sidewall and thus was relatively thick, 0.49 arcsec, and suited for pitot probe surveys. The test conditions were a nominal Mach number of 3 and unit Reynolds number of 1.2 x 10 exp 7/ft. The measurements obtained included surface oil flow visualizations, surface static pressures, and boundary layer pitot pressure profiles. The results showed that downstream of the crossing shock intersection, the stagnation pressure losses were significant and the stagnation pressure profiles were highly nonuniform. Despite the severe shock disturbances, the law of the wall and the law of the wake were found to give relatively good agreement with the experimental data.

Lachowicz, Jason T.↗