Flight investigation of the variation of static-pressure error of a static-pressure tube with distance ahead of a wing and a fuselage
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Description of a Pressure Measurement Technique for Obtaining Surface Static Pressures of a Radial Turbine
The aerodynamic performance of a solid uncooled version of a cooled radial turbine was evaluated in the Small Engine Components Test Facility Turbine rig at the NASA Lewis Research Center. Specifically, an experiment was conducted to rotor surface static pressures. This was the first time surface static pressures had been measured on a radial turbine at NASA Lewis. These pressures were measured by a modified Rotating Data Package (RDP), a standard product manufactured by Scanivalve, Inc. Described here are the RDP, and the modifications that were made, as well as the checkout, installation, and testing procedures. The data presented are compared to analytical results obtained from NASA's MERIDL TSONIC BLAYER (MTSB) code.
Dicicco, L. Danielle↗
Description of a pressure measurement technique for obtaining surface static pressures of a radial turbine
The aerodynamic performance of a solid uncooled version of a cooled radial turbine was evaluated in the Small Engine Components Test Facility Turbine rig at the NASA Lewis Research Center. Specifically, an experiment was conducted to rotor surface static pressures. This was the first time surface static pressures had been measured on a radial turbine at NASA Lewis. These pressures were measured by a modified Rotating Data Package (RDP), a standard product manufactured by Scanivalve, Inc. Described here are the RDP, and the modifications that were made, as well as the checkout, installation, and testing procedures. The data presented are compared to analytical results obtained from NASA's MERIDL TSONIC BLAYER (MTSB) code.
Dicicco, L. D.↗
Water cooled static pressure probe
An improved static pressure probe containing a water cooling mechanism is disclosed. This probe has a hollow interior containing a central coolant tube and multiple individual pressure measurement tubes connected to holes placed on the exterior. Coolant from the central tube symmetrically immerses the interior of the probe, allowing it to sustain high temperature (in the region of 2500 F) supersonic jet flow indefinitely, while still recording accurate pressure data. The coolant exits the probe body by way of a reservoir attached to the aft of the probe. The pressure measurement tubes are joined to a single, larger manifold in the reservoir. This manifold is attached to a pressure transducer that records the average static pressure.
Lagen, Nicholas T.↗
Hot-film static-pressure probe for surveying flow fields
A static pressure probe employing hot-film sensors has been developed for the rapid measurement of the static pressure fields surrounding analytic shapes in hypersonic flows. The hot-film probe is a modification of the standard static pressure probe, consisting of a front hot-film sensor operated as a resistance thermometer, a rear sensor operated at an overheat ratio of 1.5 to 1.8 and a small sonic orifice installed inside the tubing of a conventional device. The probe has been calibrated in helium and air over a range of temperatures and pressures in a bell jar apparatus, with a repeatability of the data to within + or - 0.015 mm Hg. Comparative tests of the hot-film and conventional static pressure probes in a hypersonic helium wind tunnel at Mach 20 and various Reynolds numbers have indicated the settling time of the hot-film probe to be on the order of milliseconds, as compared with 30 sec for the conventional probe. The pressures measured by the two probes were found to be within 10% of each other. Although the hot-film probe makes flow-field static pressure surveys more practical in blowdown hypersonic wind tunnels, viscous and flow angle effects still must be assessed under the conditions of use.
Ashby, G. C., Jr.↗
Measurement of static pressure on aircraft
Existing data on the errors involved in the measurement of static pressure by means of static-pressure tubes and fuselage vents are presented. The errors associated with the various design features of static-pressure tubes are discussed for the condition of zero angle of attack and for the case where the tube is inclined to flow. Errors which result from variations in the configuration of static-pressure vents are also presented. Errors due to the position of a static-pressure tube in the flow field of the airplane are given for locations ahead of the fuselage nose, ahead of the wing tip, and ahead of the vertical tail fin. The errors of static-pressure vents on the fuselage of an airplane are also presented. Various methods of calibrating static-pressure installations in flight are briefly discussed.
Gracey, William↗
Experimental cavity pressure measurements at subsonic and transonic speeds. Static-pressure results
An experimental investigation was conducted to determine cavity flow-characteristics at subsonic and transonic speeds. A rectangular box cavity was tested in the Langley 8-Foot Transonic Pressure Tunnel at Mach numbers from 0.20 to 0.95 at a unit Reynolds number of approximately 3 x 10(exp 6) per foot. The boundary layer approaching the cavity was turbulent. Cavities were tested over a range of length-to-depth ratios (l/h) of 1 to 17.5 for cavity width-to-depth ratios of 1, 4, 8, and 16. Fluctuating- and static-pressure data in the cavity were obtained; however, only static-pressure data is analyzed. The boundaries between the flow regimes based on cavity length-to-depth ratio were determined. The change to transitional flow from open flow occurs at l/h at approximately 6-8 however, the change from transitional- to closed-cavity flow occurred over a wide range of l/h and was dependent on Mach number and cavity configuration. The change from closed to open flow as found to occur gradually. The effect of changing cavity dimensions showed that if the vlaue of l/h was kept fixed but the cavity width was decreased or cavity height was increased, the cavity pressure distribution tended more toward a more closed flow distribution.
Plentovich, E. B.↗
Effects of inlet distortion on a static pressure probe mounted on the engine hub in an F-15 airplane
An inlet static pressure (PS2) probe was mounted on the hub of an F100 engine in an F-15 airplane. Flight test results showed that for low distortion conditions, the ratio of engine-face total pressure to static pressure agreed well with previous altitude facility data. Off-schedule operation of the inlet third ramp angle caused increased distortion of the inlet airflow during steady-state flight conditions. Data are shown for inlet third ramp excursions leading to engine stall. The relationships of inlet face total to static pressure ratio as a function of several distortion descriptors are also described.
Hughes, D. L.↗
Static Pressure Orifice Imaging & Data Correlation
The quality of static pressure ports has historically been shown to be capable of biasing the resulting measurements of static pressure over the surface. Some of the flaws in the port quality can be recognized visually through magnified imaging of the port with a microscope. Through digital analyses of images of a large quantity of existing pressure ports and estimation of the deviations from the true static pressure for each port, a correlation could potentially be constructed to estimate whether a port will generate a significant error based upon its surface-level appearance. This presentation pursues this hypothesis through attempted correlation of subsonic data collected on the 4-inch-diameter cone cylinder during the 2019 characterization tests in the NASA Glenn Research Center's 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT).
8x6 SWT↗
A short static-pressure probe design for supersonic flow
A static-pressure probe design concept was developed which has the static holes located close to the probe tip and is relatively insensitive to probe angle of attack and circumferential static hole location. Probes were constructed with 10 and 20 deg half-angle cone tips followed by a tangent conic curve section and a tangent cone section of 2, 3, or 3.5 deg, and were tested at Mach numbers of 2.5 and 4.0 and angles of attack up to 12 deg. Experimental results indicate that for stream Mach numbers of 2.5 and 4.0 and probe angle of attack within + or - 10 deg, values of stream static pressure can be determined from probe calibration to within about + or - 4 percent. If the probe is aligned within about 7 deg of the flow experimental results indicated, the stream static pressures can be determined to within 2 percent from probe calibration.
Pinckney, S. Z.↗
Subsonic wind-tunnel tests of a trailing-cone device for calibrating aircraft static pressure systems
A trailing-cone device for calibrating aircraft static-pressure systems was tested in a transonic wind tunnel to investigate the pressure-sensing characteristics of the device including effects of several configuration changes. The tests were conducted at Mach numbers from 0.30 to 0.95 with Reynolds numbers from (0.9 x one million to 4.1 x one million per foot). The results of these tests indicated that the pressures sensed by the device changed slightly but consistently as the distance between the device pressure orifices and cone was varied from 4 to 10 cone diameters. Differences between such device-indicated pressures and free-stream static pressure were small, however, and corresponded to Mach number differences of less than 0.001 for device configurations with pressure orifices located 5 or 6 cone diameters ahead of the cone. Differences between device-indicated and free-stream static pressures were not greatly influenced by a protection skid at the downstream end of the pressure tube of the device nor by a 2-to-1 change in test Reynolds number.
Jordan, F. L., Jr.↗
A bleed-type pressure transducer for in-stream measurement of static pressure fluctuations
Turbulent incompressible air flow in-stream static pressure fluctuations measurement, describing bleed type pressure transducer theory, design and operational characteristics
Jones, B. G.↗
Wind-Tunnel Tests of Seven Static-Pressure Probes at Transonic Speeds
Wind-tunnel tests have been conducted to determine the errors of 3 seven static-pressure probes mounted very close to the nose of a body of revolution simulating a missile forebody. The tests were conducted at Mach numbers from 0.80 to 1.08 and at angles of attack from -1.7 deg to 8.4 deg. The test Reynolds number per foot varied from 3.35 x 10(exp 6) to 4.05 x 10(exp 6). For three 4-vane, gimbaled probes, the static-pressure errors remained constant throughout the test angle-of-attack range for all Mach numbers except 1.02. For two single-vane, self-rotating probes having two orifices at +/-37.5 deg. from the plane of symmetry on the lower surface of the probe body, the static-pressure error varied as much as 1.5 percent of free-stream static pressure through the test angle-of- attack range for all Mach numbers. For two fixed, cone-cylinder probes of short length and large diameter, the static-pressure error varied over the test angle-of-attack range at constant Mach numbers as much as 8 to 10 percent of free-stream static pressure.
Capone, Francis J.↗
Static pressure orifice system testing method and apparatus
A method and apparatus are presented for pressure testing the static pressure orifices and associated connections used in wind tunnels. A cylindrical module, having in one end an open hemispherical calibration pressure chamber separated from and surrounded by an annular vacuum chamber is placed over the orifice of the system to be tested. O-rings ensure seating and a vacuum seal between the chambered end of the module and the surface around the orifice: one O-ring separates the outer chamber from the outside environment. Ports lead from each of the chambers out the other end of the module to tubes connected to a control box consisting of calibration pressure and vacuum supply lines, bleeder valves, and gauges.
Culotta, R. F.↗
A Simple Boundary-Layer Transition Detection Method Using Mean Static Pressure Measurements
A simple method for detecting boundary-layer transition using only mean static pressure port data is presented. The method can be applied to most existing models with pressure taps, and only requires that a fine angle-of-attack sweep be performed. A small but abrupt change in the static pressure is visible when the transition front passes over the pressure tap. Results from a recent Juncture Flow test entry are used to illustrate the technique. Infrared thermography measurements of the transition front compare very well to the transition locations obtained from the static pressure ports. Some differences in behavior occur depending on the dominant transition mechanism. While the technique is somewhat qualitative, it can be an excellent tool for estimating the transition location when other tools are not readily available. Given the simplicity of the technique, and the fact that most wind-tunnel models are already designed with numerous static pressure taps, this method can be applied with very little overhead.
boundary-layer transition↗
Static pressure measurements in A 30 kWe class arcjet
A series of cold and hot flow static pressure measurements have been made as a function of distance along the principal axis of a 30 kWe arcjet. During these experiments, the background pressure was varied from atmospheric pressure to less than 1 Torr. The experimental arcjet used in these tests had the same constrictor and conical nozzle dimensions as the baseline MOD-1 ammonia arcjet developed at NASA-JPL, but used nitrogen as the propellant gas. Comparison of the measured static pressure profiles with the arc erosion patterns on the anode surface showed that anode arc attachment points directly coincided with the minimum static pressure points. The dual voltage modes commonly observed in 30 kWe class arcjet thrusters are also believed to be caused by the presence of two pressure minima in the arcjet, one at the constrictor entrance and the other in the expansion nozzle. The static pressure profiles also suggest that the sonic flow transition occurs at the entrance to the nozzle, indicating the existence of very thick boundary layers in the arcjet constrictor.
Harris, William J.↗
External Interference Effects of Flow Through Static-Pressure Orifices of an Airspeed Head at Several Supersonic Mach Numbers and Angles of Attack
Wind-tunnel tests have been made to determine the static-pressure error resulting from external interference effects of flow through the static-pressure orifices of an NACA airspeed head at Mach numbers of 2.4, 3.0, and 4.0 for angles of attack of 0 deg, 5 deg, 10 deg, and 15 deg. Within the accuracy of the measurements and for the range of mass flow covered, the static-pressure error increased linearly with increasing mass-flow rate for both the forward and rear sets of orifices at all Mach numbers and angles of attack of the investigation. For a given value of flow coefficient, the static-pressure error varied appreciably with Mach number but only slightly with angle of attack. For example, for a flow coefficient out of the orifices of 0.01 (the approximate value for a vertically climbing airplane for which the airspeed system incorporates an airspeed meter, a Mach meter, and an altimeter), the error increased from about 5 percent to about 12 percent of the static pressure as the Mach number increased from 2.4 to 4.0 with the airspeed head at an angle of attack of 0 deg.
Silsby, Norman S.↗
Transonic Wind-Tunnel Tests of an Error-Compensated Static-Pressure Probe
An investigation of the pressure-sensing characteristics of an error-compensated static-pressure probe mounted on the nose section of a missile body has been conducted in the Langley 16-foot transonic tunnel. The probe was free to rotate about its roll axis and was equipped with a vane so that the crossflow velocity component due to angles of attack or sideslip was always alined with the probe's vertical plane of symmetry. The probe was tested in five axial positions with respect to the missile nose at Mach numbers from 0.30 to 1.08 and at angles of attack from -2.7 to 15.3 deg. The test Reynolds number per foot varied from 1.79 x 10(exp 6) to 4.05 x 10(exp 6). Results showed that at a Mach number of 1.00 the static-pressure error decreased from 3.5 percent to 0.8 percent of the free-stream static pressure, as a result of a change in orifice location from 0.15 maximum missile diameter to 0.20 maximum missile diameter forward of the missile nose. Although compensation for pressure-sensing errors due to angles of attack up to 15.3 was maintained at Mach numbers from M = 0.30 to M = 0.50, there was an increase in error with an increase in angle of attack for Mach numbers between M 0.50 and M = 1.08.
Capone, Francis J.↗