scholarly journals Prediction of total pressure characteristics in the settling chamber of a supersonic blowdown wind tunnel

2011 ◽  
Vol 115 (1171) ◽  
pp. 557-566 ◽  
Author(s):  
G. K. Suryanarayana ◽  
S. R. Bhoi
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Storage Tank ◽  
Stagnation Pressure ◽  
Rate Of Change ◽  
Settling Chamber ◽  
Low Mach Number ◽  
High Mach ◽  
And Storage ◽  
Pressure Characteristics

Abstract Occurrence of transient starting and stopping loads during tests at high Mach numbers is one of the major problems in intermittent blowdown wind tunnels. It is believed that in order to overcome this problem, the wind tunnel could be started at a low Mach number and low stagnation pressure; the desired high Mach number condition could be reached by continuously changing the nozzle contour while synchronously increasing the stagnation pressure. After completing the tests, the nozzle could be brought back to the initial low Mach number accompanied by synchronous decrease in the stagnation pressure. In such a scenario, it is important to ensure that the pressure regulating valve (PRV) of the wind tunnel delivers and maintains a specified minimum stagnation pressure at any Mach number, so that supersonic breakdown of the test section flow does not occur. In this paper, the problem is formulated based on quasi-steady one-dimensional isentropic equations and numerically solved to predict the time histories of settling chamber pressure and storage tank pressure for a given trajectory of the opening of the PRV, as the Mach number is changed from Mach 1 to 4·0 continuously in four seconds and vice versa. The effects of rate of change of PRV open area and rate of change of Mach number on the stagnation pressure characteristics in the settling chamber and storage tank are predicted. The measured trajectories of the PRV in experiments in the NAL 0·6m transonic wind tunnel are used as input to the prediction program to validate the methodology. Predictions indicate that when the nozzle throat is changed from Mach 1 to 4 in four seconds, the settling chamber stagnation pressure rapidly builds up and approaches the pressure in the storage tank. Predictions show an alarming rise in free stream dynamic pressure during transition from Mach 1 to 4 and vice versa, which needs to be verified through measurements.

Design Point Efficiency of Radial Turbines

2018 ◽  
Author(s):  
Graham D. Cox
Keyword(s):  
Mach Number ◽  
Numerical Solutions ◽  
Flow Coefficient ◽  
Speed Ratio ◽  
Low Mach Number ◽  
High Mach Number ◽  
Specific Speed ◽  
High Mach ◽  
Nozzle Guide ◽  
New Applications

Results are presented from CFD calculations on a large database of 3D, radially-stacked, radial turbine geometries. The database covers a comprehensive range of basic geometrical features allowing the most appropriate geometry to be selected for optimum efficiency over extensive ranges of blade-speed-ratio, flow coefficient and specific speed. Initial studies considered the wheel only and used 78 geometries for low Mach number applications and 102 geometries for high Mach number applications. Each was run over a range of blade-speed-ratios and inlet flow angles to generate preliminary results. Designs that did not contribute to the optimum efficiency trends were discarded. The remaining 17 low Mach number and 24 High Mach number wheels were recalculated with a range of nozzle guide vanes and back-face cavities to provide increased fidelity numerical solutions. The calculated optimum efficiency is presented on charts of blade-speed-ratio against flow coefficient, against specific speed and against non-dimensional mass-flow. The effect of exducer trim reduction, as often required for mechanical reasons, is demonstrated. The charts can be used for preliminary design of new applications.

Performance of a jaws inlet under off-design conditions

2016 ◽  
Vol 231 (2) ◽  
pp. 294-305 ◽  
Author(s):  
Tianlai Gu ◽  
Shuai Zhang ◽  
Yao Zheng
Keyword(s):  
Mach Number ◽  
Flow Fields ◽  
Back Pressure ◽  
Low Mach Number ◽  
Mass Capture ◽  
Constant Area ◽  
Flow Capture ◽  
Mach Numbers ◽  
High Mach ◽  
Better Than

Numerical analysis was conducted of a jaws inlet under different working conditions, including angles of attack of 0° and 3°, varying Mach number, and varying back pressure with a constant-area isolator, to investigate its performance and flow fields of starting and unstarting states. Results reveal that the jaws inlet has an enhanced flow capture capability in starting states, with the mass capture ratio higher than 0.96, but relatively reduced working range of inflow Mach numbers. Its performance at a low Mach number is better than that at a high Mach number. Non-uniform flow fields are observed in unstarting cases at low Mach numbers and high back pressures, while separation structures are confined in the pitching compression section. Further increase in Mach number or decrease in back pressure does not result in significant changes in the separation structures. In the unstarting case under a high back pressure, it is hard to achieve restarting through reductions in the back pressure.

The Ludwieg Pressure-Tube Supersonic Wind Tunnel

1963 ◽  
Vol 14 (2) ◽  
pp. 143-157 ◽  
Author(s):  
A. J. Cable ◽  
R. N. Cox
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Nozzle Exit ◽  
High Speed ◽  
Layer Growth ◽  
Stagnation Pressure ◽  
Pressure Tube ◽  
Rarefaction Waves ◽  
Supersonic Wind Tunnel ◽  
Blow Down

SummaryA description is given of a supersonic pressure-tube wind tunnel which has been constructed at A.R.D.E. This is a blow-down tunnel which uses as a reservoir a long tube filled with gas under pressure. A quasi-steady supersonic flow is achieved by expanding in a convergent-divergent nozzle the subsonic flow behind rarefaction waves which propagate down the tube when a diaphragm at the nozzle exit is burst. The theory of the operation of the tunnel is given and calculations are made of the boundary-layer growth along the tube. Pressure-time records were obtained in the tube, and a high speed camera was used to obtain pictures of the flow round a model. Measurements also included a pitot-tube traverse of the nozzle exit, and the Mach number distribution was determined from the ratio of the pitot to the stagnation pressure. Tests showed that, as predicted, a constant stagnation pressure was obtained ahead of the nozzle, and it is considered that a tunnel of this type would be a cheap and simple way of obtaining an intermittent tunnel with adequate running time for many types of test, and capable of operating at a Reynolds number of more than 107 per inch at a Mach number of about 3·5.

scholarly journals Aspects of aero-engine nacelle drag

2018 ◽  
Vol 233 (5) ◽  
pp. 1667-1682 ◽  
Author(s):  
Matthew Robinson ◽  
David G MacManus ◽  
Christopher Sheaf
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Mach Number ◽  
Wind Tunnel ◽  
Near Field ◽  
Pressure Ratio ◽  
Exhaust System ◽  
Extraction Methods ◽  
Field Method ◽  
Rate Of Change

To address the need for accurate nacelle drag estimation, an assessment has been made of different nacelle configurations used for drag evaluation. These include a sting mounted nacelle, a nacelle in free flow with an idealised, freestream pressure matched, efflux and a nacelle with a full exhaust system and representative nozzle pressure ratio. An aerodynamic analysis using numerical methods has been carried out on four nacelles to assess a near field drag extraction method using computational fluid dynamics. The nacelles were modelled at a range of aerodynamic conditions and three were compared against wind tunnel data. A comparison is made between the drag extraction methods used in the wind tunnel analysis and the chosen computational fluid dynamics approach which utilised the modified near-field method for evaluation of drag coefficients and trends with Mach number and mass flow. The effect of sting mounting is quantified and its influence on the drag measured by the wind tunnel methodology determined. This highlights notable differences in the rate of change of drag with free stream Mach number, and also the flow over the nacelle. A post exit stream tube was also found to create a large additional interference term acting on the nacelle. This term typically accounts for 50% of the modified nacelle drag and its inclusion increased the drag rise Mach number by around Δ M = 0.026 from [Formula: see text] to [Formula: see text] for the examples considered.

Experimental and numerical investigation of a supersonic inlet with double S-bend diffuser

2021 ◽  
pp. 095441002098398
Author(s):  
Jinsheng Zhang ◽  
Huacheng Yuan ◽  
Yunfei Wang ◽  
Guoping Huang
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Shock Train ◽  
Recovery Coefficient ◽  
Low Mach Number ◽  
Supersonic Inlet ◽  
Total Pressure Recovery ◽  
High Mach ◽  
Pressure Recovery Coefficient

Design of a supersonic inlet with double S-bend diffuser was developed. Numerical simulation and wind tunnel experiment were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the variable geometry scheme adopted solves the contradiction between starting performance at low Mach number and aerodynamic performance at high Mach number. The inlet works normally and stably over a wide speed range. At design point, the total pressure recovery coefficient reaches 0.47. In addition, two different kinds of inlets with double S-bend diffuser and single S-bend diffuser were studied. Compared with the double S-bend diffuser, the total pressure recovery coefficient of the single S-bend diffuser is higher at low Mach number (Ma0 < 3) and lower at high Mach number (Ma0 > 3). With the increase of backpressure, shock train mainly moves upstream along the low-energy flow region in the diffuser. For the double S-bend diffuser, shock train will first move along the lower corner and then along the upper corner. For the single S-bend diffuser, it will only move along the upper corner. The strong secondary flow of the double S-bend is the main reason for the above phenomenon.

scholarly journals Conversion of an Aerodynamic Wind Tunnel to a Gas Turbine Combustor Test Stand

1995 ◽  
Author(s):  
Daniel Cresci
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Pressure Control ◽  
Test Facility ◽  
Gas Turbine Combustor ◽  
Low Mach Number ◽  
Air Supply ◽  
Emission Testing ◽  
Aerodynamic Test ◽  
High Mach

A Mach 12, blowdown-type, storage heated, aerodynamic test facility, previously adapted to provide Mach 3.5 ramjet engine propulsion testing, has been further modified to provide high pressure, gas turbine combustor emission testing capability. Pressure control and temperature control systems have been incorporated into the original facility to provide constant flow properties for run times up to 30 minutes. An on-line gas sample system has also been designed and integrated with the facility to provide continuous combustor emission information. These features now combine with the existing pebble bed storage heater, blowdown air supply and heated fuel supply systems to provide very high Mach number aerodynamic simulation, intermediate Mach number aeropropulsion testing and low Mach number, full scale, gas turbine combustor emission testing capability. Pressures of 1500 psia can be achieved with temperatures up to 2000 degrees Fahrenheit and flow rates up to 50 lb/sec. This unique facility is described, with emphasis on the facility modifications necessary to provide the gas turbine combustor emission testing capability. To appreciate the modifications made, some of the differences between Mach 12 aerodynamic and low Mach number, gas turbine combustor emission testing are discussed. Results from some recent gas turbine emissions tests are also provided.

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