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106 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Internal performance of a nonaxisymmetric nozzle with a rotating upper flap and a center-pivoted lower flap

An investigation was conducted at wind-off conditions in the static-test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance characteristics of a single expansion-ramp nozzle with thrust-vectoring capability to 105 degrees. Thrust vectoring was accomplished by the downward rotation of an upper flap with adaptive capability for internal contouring and a corresponding rotation of a center-pivoted lower flap. The static internal performance of configurations with pitch thrust-vector angles of 0 degrees, 60 degrees, and 105 degrees each with two throat areas, was investigated. The nozzle pressure ratio was varied from 1.5 to approximately 8.0 (5.0 for the maximum throat area configurations). Results of this study indicated that the nozzle configuration of the present investigation, when vectored, provided excellent flow-turning capability with relatively high levels of internal performance. In all cases, the thrust vector angle was a function of the nozzle pressure ratio. This result is expected because the flow is bounded by a single expansion surface on both vectored- and unvectored-nozzle geometries.

Wing, David J.↗
NASA NTRS1980

Low speed test of the aft inlet designed for a tandem fan V/STOL nacelle

An approximately .25 scale model of a Tandem Fan nacelle designed for a Type A V/STOL aircraft configuration was tested in a 10-by-10 foot wind tunnel. A 12 inch, tip driven, turbofan simulator was used to provide the suction source for the aft fan inlet. The front fan inlet was faired over for this test entry. Model variables consisted of a long aft inlet cowl, a short aft inlet cowl, a shaft simulator, blow-in door passages and diffuser vortex generators. Inlet pressure recovery, distortion, inlet angle of attack separation limits were evaluated at tunnel velocities from 0 to 240 knots, angles of attack from -10 to 40 degrees and inlet flow rates representative of throat Mach numbers of 0.1 to 0.6. High inlet performance and stable operation was verified at all design forward speed and angle of attack conditions. The short aft inlet configuration provided exceptionally high pressure recovery except at the highest combination of angle of attack and forward speed. The flow quality at the fan face was somewhat degraded by the addition of blow-in door passages to the long aft inlet configuration due to the pressure disturbances generated by the flow entering the diffuser through the auxiliary air passages.

Rhoades, W. W.↗
NASA NTRSDate not supplied

Testing of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future Mars human lander missions will require using retrorockets beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes, the freestream flow, and the vehicle that alter the aerodynamic behavior. Historical retropropulsion ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. A new retropropulsion wind tunnel test was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to address some of the deficiencies of previous tests and to provide a basis for computational model comparisons on relevant geometries. The test included improvements in model design and measured data. Two sub-scale model geometries derived from full-scale vehicle concepts were tested: a blunt low lift-to-drag vehicle and a more slender geometry. The first model includes variations in nozzle parameters that were estimated to have first-order effects on the measurements: location, pointing direction, size, and exit-to-throat area ratio. Both models were instrumented with discrete steady and unsteady pressure sensors on the aerodynamic and interior flow path surfaces. The heats shields of both models also were covered with pressure sensitive paint that effectively measures a steady pressure distribution in an otherwise unsteady flow field. This pressure distribution was integrated over the surface geometry in order to estimate the steady aerodynamic force and moment coefficient contribution from the painted surfaces; this approach was used as a replacement for using a flow-through balance. This paper will cover the test objectives, model design, test matrix, and available test results for each model configuration at different freestream Mach numbers, model attitude, and non-dimensional thrust magnitude. A subsequent publication will cover uncertainty quantification of the test data and comparisons to computational fluid dynamics model results.

Powered Descent↗
NASA NTRS1992

Mach 5 electroformed nickel nozzle refurbishment FNAS investigation of ultra-smooth surfaces

This task is in support of the Quiet Hypersonic Wind Tunnel effort currently in effect at NASA Langley Research Center, VA. A laminar flow wind tunnel nozzle has been previously fabricated by electroforming pure nickel over a machined and polished two piece stainless steel mandrel. The mandrel was removed leaving the replicate nozzle surface. The nozzle was then pressed into a heavy stainless steel jacket for mounting features and rigidity. The original nickel surface was a replication of the polished mandrel but had degraded due to oxidation. The inside surface requirements are very stringent in order to achieve laminar or quiet flow at the desired pressure and temperature for the specific design of Mach 5. The throat area of the axisymmetric device must have a surface finish with no defects greater than 16 microinches. This requires an rms average background of about four microinches or better for inspection purposes. A coating of nickel-phosphorous alloy was applied by catalytic deposition and then the inside of the nozzle was polished retaining dimensional and surface finish tolerances as specified per drawings supplied. Since the unit is not an optical component, conventional optical inspection methods for surface finish and figure are not readily achieved. Measurements were made using surface profilometry.

Rood, Robert↗
NASA NTRS1953

Preliminary Evaluation of Turbine Performance with Variable-Area Turbine Nozzles in a Turbojet Engine

The performance of a two-stage turbine with variable-area first-stage turbine nozzles was determined in the NACA Lewis altitude wind tunnel over a range of simulated altitudes from 15,000 to 44,000 feet and engine speeds from 50 to 100 percent of rated speed. The variable-area turbine nozzles used in this investigation were primarily a test device for compressor research purposes and were not necessarily of optimum aerodynamic design. The results of this investigation are indicative of effects of turbine-nozzle-area variation on turbine performance within the operating range allowed by the engine. The variable-area turbine nozzles were found to be mechanically reliable and to have negligible leakage losses. Increasing the turbine-nozzle-throat area from 1.15 to 1.67 square feet increased the corrected turbine gas flow or effective turbine nozzle area about 10 percent. At a given corrected turbine speed and turbine pressure ratio, changing the turbine nozzle area from 1.30 to 1. 67 square feet lowered the turbine efficiency 3 or 4 percent. The effect of increasing the turbine nozzle area from 1.15 to 1.67 square feet (decreasing the turning angle about 7 1/2 degrees) would be to lower the turbine efficiency about 5 or 6 percent.

Campbell, Carl E.↗
NASA NTRS1981

Additional testing of the inlets designed for a tandem fan V/STOL nacelle

The wind tunnel testing of a scale model of a tandem fan nacelle designed for a type (subsonic cruise) V/STOL aircraft configuration is discussed. The performance for the isolated front inlet and for the combined front and aft inlets is reported. Model variables include front and aft inlets with aft inlet variations of short and long aft inlet cowls, with a shaft simulator and diffuser vortex generators, cowl lip fillets, and nacelle strakes. Inlet pressure recovery, distortion, and inlet angle-to-attack separation limits were evaluated at tunnel velocity from 0 to 240 knots, angles-of-attack from -10 to +40 degrees and inlet flow rates corresponding to throat Mach number from 0.0 to 0.6. Combined nacelle pitch and yaw runs up to 30 deg. were also made.

Ybarra, A. H.↗
NASA NTRS1981

Zero-length inlets for subsonic V/STOL aircraft

Zero-length inlet performance and associated fan blade stresses were determined during model tests in the NASA-LeRC 9-by 15-foot low-speed wind tunnel. The inlet models, which were installed on a 20-inch diameter fan unit, had different inlet lip contraction ratios as well as unslotted, slotted, and double slotted inlet lips. The inlet angle-of-attack boundaries for onset of flow separation were identified and compared to the operating requirements of several generically different subsonic V/STOL aircraft. The zero-length inlets, especially those with slotted lips, were able to satisfy these requirements without compromising the maximum cowl forebody radius. As an aid to the inlet design process, a unique relationship was established between the maximum surface Mach number associated with the separation boundary and the maximum-to-throat surface velocity ratio.

Glasgow, E. R.↗
NASA NTRS1976

Inlets for high angles of attack

Different inlet designs for high angle of attack STOL and VTOL applications were tested in a subsonic wind tunnel. Three removable entry lips having contraction ratios of 1.30, 1.34 and 1.38 were tested with a single diffuser. The internal contour of each entry lip was an ellipse with a major to minor axis of 2.0. Each lip and diffuser assembly was tested to determine its tolerance to angle of attack, first with a conventional centerbody and then with an extended centerbody. Results indicate that a large improvement in separation angle (determined as a function of lip contraction ratio and inlet flow) was obtained for the extended centerbody for all contraction ratios. Improved inlet tolerance to angle of attack was obtained by reducing the adverse pressure gradient downstream of the throat.

Miller, B. A.↗
NASA NTRSDate not supplied

Seeding Method for Velocimetry and Visualization of Supersonic Retropropulsion Nozzle Plumes

In the current work, the “Venturi seeder” method for nozzle plumes is improved and studied to determine suitability for particle image velocimetry. This seeding method involves independently pressurizing the primary flow path and a separate seed liquid reservoir, connecting them at the throat of a Venturi contraction upstream of the nozzle. The pressure differential at the throat causes the liquid to enter the nozzle channel where it is atomized into particles by strong shearing forces. This new seeding system was characterized by conducting a series of diagnostics using different exit nozzle pressure ratios and seed reservoir pressures. Shadowgraph imaging confirmed the nozzle plumes to be underexpanded jets. A particle sizing device determined that generated DEHS oil particles had median aerodynamic diameters of 0.67, 0.69, and 0.73 μm for seed liquid reservoir pressure supplies of 0.67, 0.60, and 0.93 MPa and primary flow path pressures of 0.67, 0.76, and 0.93 MPa measured downstream of the Venturi throat, respectively. These diameters were within the threshold for acceptable response time in typical gas flows (≤ 1 μm). PIV experiments were conducted on the nozzle plume for all presented cases; mean axial and transverse velocities appeared as expected for the underexpanded jet structure, including the Mach disk and re-acceleration regions. Additional sizing analysis based on the particle response times across the normal shock again showed particle diameters to be well less than 1 μm for all cases. The results suggest that this method is an inexpensive and relatively simple solution to the problem of seeding nozzle plumes. The method will be used in testing of supersonic retropropulsion models in the Unitary Plan Wind Tunnel at The NASA Langley Research Center in 2022.

particle image velocimetry↗
NASA NTRS2023

Passive Rocket Diffuser Testing: Reacting Flow Performance of Advanced Configurations

Upper stage rocket engines are optimized for operation at the edge of space. Ground testing often requires simulation of high-altitude environments to prevent damage from off-design operation. Passive supersonic diffusers are an effective means of establishing the partial vacuum needed around the nozzle exit. The low cost of diffusers relative to active vacuum sources (ejectors, pumps, etc.) provides a perpetual incentive to improve their performance and expand the envelope of passively testable engines. Second-throat diffusers have been the de facto standard for rocket testing since the 1950s, and numerous studies have focused on parametric optimization of the topology. Yet, even the best second-throat designs may require more driving pressure than an upper stage engine can provide. To overcome conventional limits, engineers at NASA’s Stennis Space Center (SSC) adapted an obscure cold-flow diffuser topology to the extremes of immersion in supersonic combustion products. Aerodynamic experimentation was conducted at SSC’s E-3 test stand to quantify the performance of prototypical hot-fire spike diffuser hardware and enable direct comparison to prior states of the art. A LO X /GH 2 thruster was contoured to approximate the regeneratively-cooled nozzle of an upper-stage engine and fired into 28 downstream diffuser configurations: 16 second-throat, 2 centerbody, and 10 spike. Temperature and pressure were recorded along the test article walls for a range of chamber pressures spanning 3.5 - 5.0 MPa. The second-throat configurations with the lowest start and unstart pressure ratios were shown to outperform equivalent systems in literature and taken as baselines for evaluation of advanced diffuser performance. Compared to these exemplars, centerbody diffusers decreased start pressure ratio by 11% but increased unstart pressure ratio by 15%. Spike diffusers reduced start and unstart pressure ratios by 24% and 9%, respectively.

Rocket↗
NASA NTRS1978

Wind tunnel performance tests of coannular plug nozzles

Wind tunnel performance test results and data analyses are presented for dual-flow plug nozzles applicable to supersonic cruise aircraft during takeoff and low-speed flight operation. Outer exhaust stream pressure ratios from 1.5 to 3.5 were tested; inner exhaust stream conditions were varied from very low, or bleed flow rates, up to a pressure ratio of 3.5. Mach numbers tested ranged from zero to 0.45. Measured thrust coefficients for the eight model configurations, operating at an external Mach number of 0.36 and an outer flow pressure ratio of 2.5, varied from 0.95 to 0.974 for high inner flow rates. At low inner flow, the performance ranged from 0.88 to 0.97 for the same operating conditions. The primary design variables influencing the performance levels were the annular height of the inner and outer nozzle throats (denoted by radius ratio - the ratio of inner-to-outer flowpath diameter at the nozzle throat), the plug geometry, and the inner stream flow rate.

Staid, P. S.↗
NASA NTRS1975

Experimental evaluation of a Mach 3.5 axisymmetric inlet

Wind tunnel test results for a large scale inlet model designed for Mach 3.5 are presented and compared with analytical predictions. The inlet is an axisymmetric mixed-compression type with a lip diameter of 49.723 cm. The inlet design was developed using analytical procedures. Data are shown for freestream Mach numbers from 0.6 to 3.5. The test results indicate that boundary layer bleed requirements can be accurately predicted. Good agreement was obtained with analytical predictions of the flowfield structure and boundary layer development in the supersonic diffuser yielding high performance at the design Mach number. The highest engine face total pressure recovery at Mach 3.5 was 85.8%; this was obtained at 0.05 Mach tolerance with only 2.8% total pressure distortion and 13.4% bleed. In the started Mach number range from 1.6 to 3.5, the total pressure recovery in the throat, downstream of the terminal normal shock, ranged between 91% and 95%. Total pressure losses in the subsonic diffuser varied from 3% to 13%. The higher losses occuring between Mach 2.5 and 3.2 were believed to be caused by the rapid rate of increase in the area of the diffuser just downstream of the throat, possibly coupled with inadequate centerbody throat bleed. In the unstarted mode at transonic speeds, the maximum inlet flow was over 99% of the theoretical maximum capture mass-flow.

Syberg, J.↗
NASA NTRS1985

Static internal performance of a single-engine onaxisymmetric-nozzle vaned-thrust-reverser design with thrust modulation capabilities

An investigation has been conducted at wind-off conditions in the stati-test facility of the Langley 16-Foot Transonic Tunnel. The tests were conducted on a single-engine reverser configuration with partial and full reverse-thrust modulation capabilities. The reverser design had four ports with equal areas. These ports were angled outboard 30 deg from the vertical impart of a splay angle to the reverse exhaust flow. This splaying of reverser flow was intended to prevent impingement of exhaust flow on empennage surfaces and to help avoid inlet reingestion of exhaust gas when the reverser is integrated into an actual airplane configuration. External vane boxes were located directly over each of the four ports to provide variation of reverser efflux angle from 140 deg to 26 deg (measured forward from the horizontal reference axis). The reverser model was tested with both a butterfly-type inner door and an internal slider door to provide area control for each individual port. In addition, main nozzle throat area and vector angle were varied to examine various methods of modulating thrust levels. Other model variables included vane box configuration (four or six vanes per box), orientation of external vane boxes with respect to internal port walls (splay angle shims), and vane box sideplates. Nozzle pressure ratio was varied from 2.0 approximately 7.0.

Leavitt, L. D.↗
NASA NTRS2001

Supersonic Wind Tunnel Tests of a Half-axisymmetric 12 Deg-spike Inlet to a Rocket-based Combined-cycle Propulsion System

Results of an isolated inlet test for NASA's GTX air-breathing launch vehicle concept are presented. The GTX is a Vertical Take-off/ Horizontal Landing reusable single-stage-to-orbit system powered by a rocket-based combined-cycle propulsion system. Tests were conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel during two entries in October 1998 and February 1999. Tests were run from Mach 2.8 to 6. Integrated performance parameters and static pressure distributions are reported. The maximum contraction ratios achieved in the tests were lower than predicted by axisymmetric Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD). At Mach 6, the maximum contraction ratio was roughly one-half of the CFD value of 16. The addition of either boundary-layer trip strips or vortex generators had a negligible effect on the maximum contraction ratio. A shock boundary-layer interaction was also evident on the end-walls that terminate the annular flowpath cross section. Cut-back end-walls, designed to reduce the boundary-layer growth upstream of the shock and minimize the interaction, also had negligible effect on the maximum contraction ratio. Both the excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls were identified as possible reasons for the discrepancy between the CFD predictions and the experiment. It is recommended that the centerbody spike and throat angles be reduced in order to lessen the induced pressure rise. The addition of a step on the cowl surface, and planar end-walls more closely approximating a plane of symmetry are also recommended. Provisions for end-wall boundary-layer bleed should be incorporated.

DeBonis, J. R.↗
NASA NTRS1993

A CFD-based aerodynamic design procedure for hypersonic wind-tunnel nozzles

A new procedure which unifies the best of current classical design practices, computational fluid dynamics (CFD), and optimization procedures is demonstrated for designing the aerodynamic lines of hypersonic wind-tunnel nozzles. The new procedure can be used to design hypersonic wind tunnel nozzles with thick boundary layers where the classical design procedure has been shown to break down. An efficient CFD code, which solves the parabolized Navier-Stokes (PNS) equations using an explicit upwind algorithm, is coupled to a least-squares (LS) optimization procedure. A LS problem is formulated to minimize the difference between the computed flow field and the objective function, consisting of the centerline Mach number distribution and the exit Mach number and flow angle profiles. The aerodynamic lines of the nozzle are defined using a cubic spline, the slopes of which are optimized with the design procedure. The advantages of the new procedure are that it allows full use of powerful CFD codes in the design process, solves an optimization problem to determine the new contour, can be used to design new nozzles or improve sections of existing nozzles, and automatically compensates the nozzle contour for viscous effects as part of the unified design procedure. The new procedure is demonstrated by designing two Mach 15, a Mach 12, and a Mach 18 helium nozzles. The flexibility of the procedure is demonstrated by designing the two Mach 15 nozzles using different constraints, the first nozzle for a fixed length and exit diameter and the second nozzle for a fixed length and throat diameter. The computed flow field for the Mach 15 least squares parabolized Navier-Stokes (LS/PNS) designed nozzle is compared with the classically designed nozzle and demonstrates a significant improvement in the flow expansion process and uniform core region.

Korte, John J.↗
NASA NTRS1951

Wind-tunnel investigation of six shielded total-pressure tubes at high angles of attack subsonic speeds

The effect of inclination of the air stream on the measured pressures of six shielded total-pressure tubes (all designed for end-mounting on a horizontal boom) has been determined for an angle-of-attack range of approximately -30° to 65°. The tests were conducted at a Mach number of 0.26 in the Langley stability tunnel and at Mach numbers of 0.50, 0.70, 0.90, and 0.95 in the Langley high-speed 7- by 10-foot tunnel. The results of the tests showed that curved venturi entries are less sensitive to inclination of the air stream than the conical entry of the standard Kiel design. Of the three curved entries tested, the tube having the bluntest entry proved the least sensitive to inclination. The critical angle (that is, the sngle.at which the total-pressure error equals 1 percent of the impact pressure) for this tube was ±63.2° at a Mach number of 0.26. The tests also showed that the critical angle of a shielded tube could be extended to higher positive angles of attack by means of a slant profile. The critical angle of a tube having a conical entry with a 10° slant profile, for example, was found to average 9° greater than that of a similar tube having a square profile. Although the critical angle of this tube at angles of yaw was less than that at angles of attack, the critical angle in yaw was equal to that of the tube with the square profile. Tests of the effect of varying the position of the total-pressure probe in the shield showed that, in general, no advantage was to be gained by changing from the position used in the standard Kiel design. Limited tests of the effect of varying the internal diameter of the shield showed that an increase in the diameter of the throat corresponding an increase in the internal area of 50 percent resulted in a relatively small increase (2.5°) in the critical angle. The effect of Mach number on the sensitivity of the tubes to inclination was shown to be appreciable. An increase in Mach number from 0.26 to 0.90, for example, resulted in a decrease in the critical angle of most of the tubes of about 4°.

Walter R Russell↗