scholarly journals Shock Characteristics Measured Upstream of Both a Forward-Swept and an Aft-Swept Fan

2007 ◽  
Author(s):  
Gary G. Podboy ◽  
Martin J. Krupar ◽  
Daniel L. Sutliff ◽  
Csaba Horvath
Keyword(s):  
Flow Visualization ◽  
Leading Edge ◽  
Suction Side ◽  
Surface Flow ◽  
Rotor Blades ◽  
Laser Doppler Velocimeter ◽  
Pressure Measurements ◽  
Pressure Transducers ◽  
Unsteady Pressure ◽  
Unsteady Pressure Measurements

Three different types of diagnostic data — blade surface flow visualization, shroud unsteady pressure, and laser Doppler velocimeter (LDV) — were obtained on two fans, one forward-swept and one aft-swept, in order to learn more about the shocks which propagate upstream of these rotors when they are operated at transonic tip speeds. Flow visualization data are presented for the forward-swept fan operating at 13831 RPMc and for the aft-swept fan operating at 12500 and 13831 RPMc (corresponding to tip rotational Mach numbers of 1.07 and 1.19, respectively). The flow visualization data identify where the shocks occur on the suction side of the rotor blades. These data show that at the takeoff speed, 13831 RPMc, the shocks occurring in the tip region of the forward-swept fan are further downstream in the blade passage than with the aft-swept fan. Shroud unsteady pressure measurements were acquired using a linear array of 15 equally-spaced pressure transducers extending from two tip axial chords upstream to 0.8 tip axial chords downstream of the static position of the tip leading edge of each rotor. Such data are presented for each fan operating at one subsonic and five transonic tip speeds. The unsteady pressure data show relatively strong detached shocks propagating upstream of the aft-swept rotor at the three lowest transonic tip speeds, and weak, oblique pressure disturbances attached to the tip of the aft-swept fan at the two highest transonic tip speeds. The unsteady pressure measurements made with the forward-swept fan do not show strong shocks propagating upstream of that rotor at any of the tested speeds. A comparison of the forward-swept and aft-swept shroud unsteady pressure measurements indicates that at any given transonic speed the pressure disturbance just upstream of the tip of the forward-swept fan is much weaker than that of the aft-swept fan. The LDV data suggest that at 12500 and 13831 RPMc, the forward-swept fan swallowed the passage shocks occurring in the tip region of the blades, whereas the aft-swept fan did not. Due to this difference, the flows just upstream of the two fans were found to be quite different at both of these transonic speeds. Nevertheless, despite distinct differences just upstream of the two rotors, the two fan flows were much more alike about one axial blade chord further upstream. As a result, the LDV data suggest that it is unwise to attempt to determine the effect that the shocks have on far field noise by focusing only on measurements (or CFD predictions) made very near the rotor. Instead, these data suggest that it is important to track the shocks throughout the inlet.

scholarly journals Effects of Transducer Installation on Unsteady Pressure Measurements on Oscillating Blades

2005 ◽  
Author(s):  
J. Lepicovsky
Keyword(s):  
Suction Side ◽  
Forced Oscillations ◽  
Pressure Measurements ◽  
Signal Distortion ◽  
Blade Surface ◽  
Pressure Transducers ◽  
Unsteady Pressure ◽  
Custom Made ◽  
Oscillation Frequencies ◽  
Unsteady Pressure Measurements

Unsteady pressures were measured above the suction side of a blade that was oscillated to simulate blade stall flutter. Measurements were made at blade oscillation frequencies up to 500 Hz. Two types of miniature pressure transducers were used: surface-mounted flat custom-made, and commercial miniature, body-mounted transducers. The signals of the surface-mounted transducers are significantly affected by blade acceleration, whereas the signals of body-mounted transducers are practically free of this distortion. A procedure was introduced to correct the signals of surface-mounted transducers to rectify the signal distortion due to blade acceleration. The signals from body-mounted transducers, and corrected signals from surface-mounted transducers represent true unsteady pressure signals on the surface of a blade subjected to forced oscillations. However, the use of body-mounted commercial transducers is preferred for the following reasons: no signal corrections are needed for blade acceleration, the conventional transducers are noticeably less expensive than custom-made flat transducers, the survival rate of body-mounted transducers is much higher, and finally installation of body-mounted transducers does not disturb the blade surface of interest.

Correcting for Response Lag in Unsteady Pressure Measurements in Water

1993 ◽  
Vol 115 (4) ◽  
pp. 676-679 ◽  
Author(s):  
Rand N. Conger ◽  
B. R. Ramaprian
Keyword(s):  
Dynamic Response ◽  
Pressure Transducer ◽  
Water Channel ◽  
Inadequate Response ◽  
Pressure Measurements ◽  
Pressure Transducers ◽  
Pitching Airfoil ◽  
Unsteady Pressure ◽  
Unsteady Pressure Measurements ◽  
Type Pressure

There is not much information available on the use of diaphragm-type pressure transducers for the measurement of unsteady pressures in liquids. A procedure for measuring the dynamic response of a pressure transducer in such applications and correcting for its inadequate response is discussed in this report. An example of the successful use of this method to determine unsteady surface pressures on a pitching airfoil in a water channel is presented.

Effect of back pressure and freestream dynamic pressure on a typical Ramjet engine duct under realistic supersonic inlet condition

2017 ◽  
Vol 122 (1247) ◽  
pp. 83-103 ◽  
Author(s):  
R. Saravanan ◽  
S.L.N. Desikan ◽  
T.M. Muruganandam
Keyword(s):  
Dynamic Pressure ◽  
Back Pressure ◽  
Flow Visualisation ◽  
Pressure Measurements ◽  
Shock Train ◽  
Unsteady Pressure ◽  
Expansion Waves ◽  
Ramjet Engine ◽  
Flap Deflection ◽  
Unsteady Pressure Measurements

ABSTRACTThe present study investigates the behaviour of the shock train in a typical Ramjet engine under the influence of shock and expansion waves at the entry of a low aspect ratio (1:0.75) rectangular duct/isolator at supersonic Mach number (M = 1.7). The start/unstart characteristics are investigated through steady/unsteady pressure measurements under different back and dynamic pressures while the shock train dynamics are captured through instantaneous Schlieren flow visualisation. Two parameters, namely pressure recovery and the pressure gradient, is derived to assess the duct/isolator performance. For a given back pressure, with maximum blockage (9% above nominal), the duct/isolator flow is established when the dynamic pressure is increased by 23.5%. The unsteady pressure measurements indicate different scales of eddies above 80 Hz (with and without flap deflection). Under the no flap deflection (no back pressure) condition, the maximum fluctuating pressure component is 0.01% and 0.1% of the stagnation pressure at X/L = 0.03 (close to the entry of the duct) and X/L = 0.53 (middle of the duct), respectively. Once the flap is deflected (δ = 8°), decay in eddies by one order is noticed. Further increase in back pressure (δ ≥ 11°) leads the flow to unstart where eddies are observed to be disappeared.

Unsteady Pressure Measurements in a Heated Rotating Cavity

2021 ◽  
Author(s):  
Richard W. Jackson ◽  
Hui Tang ◽  
James A. Scobie ◽  
Oliver J. Pountney ◽  
Carl M. Sangan ◽  
...  
Keyword(s):  
Vortex Pair ◽  
Practical Significance ◽  
Pressure Measurements ◽  
Rotating Disc ◽  
Frequency Spectra ◽  
Paper Properties ◽  
Unsteady Pressure ◽  
Rotating Cavity ◽  
Buoyancy Forces ◽  
Unsteady Pressure Measurements

Abstract The flow in the heated rotating cavity of an aero-engine compressor is driven by buoyancy forces, which result in pairs of cyclonic and anticyclonic vortices. The resultant cavity flow field is three-dimensional, unsteady and unstable, which makes it challenging to model the flow and heat transfer. In this paper, properties of the vortex structures are determined from novel unsteady pressure measurements collected on the rotating disc surface over a range of engine-representative parameters. These measurements are the first of their kind with practical significance to the engine designer and for validation of computational fluid dynamics. One cyclonic/anticyclonic vortex pair was detected over the experimental range, despite the measurement of harmonic modes in the frequency spectra at low Rossby numbers. It is shown that these modes were caused by unequal size vortices, with the cyclonic vortex the larger of the pair. The structures slipped relative to the discs at a speed typically around 10% to 15% of that of the rotor, but the speed of precession was often unsteady. The coherency, strength and slip of the vortex pair increased with the buoyancy parameter, due to the stronger buoyancy forces, but they were largely independent of the rotational Reynolds number.

scholarly journals Unsteady Pressure Measurements on a Biconvex Airfoil in a Transonic Oscillating Cascade

1986 ◽  
Vol 108 (1) ◽  
pp. 53-59 ◽  
Author(s):  
L. M. Shaw ◽  
D. R. Boldman ◽  
A. E. Buggele ◽  
D. H. Buffum
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Dynamic Pressure ◽  
Leading Edge ◽  
Pressure Transducers ◽  
High Incidence ◽  
Shock Motion ◽  
Plate Analysis ◽  
Phase Angles ◽  
Unsteady Pressure Measurements

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfoils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence angle of 0.0 deg and a Mach number of 0.65 in order to obtain results in a shock-free compressible flow field. Subsequent tests were performed at an angle of attack of 7.0 deg and a Mach number of 0.80 in order to observe the surface pressure response with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and −90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semichord reduced frequencies up to about 0.5 at a Mach number of 0.80). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

The Effect of Vanes and Blades on Ingress in Gas Turbines

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
Fabian P. Hualca ◽  
Joshua T. M. Horwood ◽  
Carl M. Sangan ◽  
Gary D. Lock ◽  
James A. Scobie
Keyword(s):  
Gas Turbines ◽  
Large Scale ◽  
Strong Influence ◽  
Rotor Blades ◽  
Pressure Measurements ◽  
Computational Results ◽  
Test Rig ◽  
Rotor Disk ◽  
Unsteady Pressure Measurements ◽  
The Relationship

Abstract This paper presents experimental and computational results using a 1.5-stage test rig designed to investigate the effects of ingress through a double radial overlap rim-seal. The effect of the vanes and blades on ingress was investigated by a series of carefully controlled experiments: first, the position of the vane relative to the rim seal was varied; second, the effect of the rotor blades was isolated using a disk with and without blades. Measurements of steady pressure in the annulus show a strong influence of the vane position. The relationship between sealing effectiveness and purge flowrate exhibited a pronounced inflection for intermediate levels of purge; the inflection did not occur for experiments with a bladeless rotor. Shifting the vane closer to the rim-seal, and therefore the blade, caused a local increase in ingress in the inflection region; again, this effect was not observed for the bladeless experiments. Unsteady pressure measurements at the periphery of the wheel-space revealed the existence of large-scale pressure structures (or instabilities) which depended weakly on the vane position and sealing flowrate. These were measured with and without the blades on the rotor disk. In all cases, these structures rotated close to the disk speed.

Experimental Investigation of the Aerodynamic Stability of an Annular Compressor Cascade Performing Tuned Pitching Oscillations in Transonic Flow

1999 ◽  
Author(s):  
H. Hennings ◽  
J. Belz
Keyword(s):  
Transonic Flow ◽  
Suction Side ◽  
Test Facility ◽  
Feedback Signal ◽  
Detailed Knowledge ◽  
Compressor Cascade ◽  
Aerodynamic Stability ◽  
Pressure Transducers ◽  
Unsteady Pressure ◽  
Unsteady Aerodynamic

A prerequisite for aeroelastic stability investigations on vibrating compressor cascades is the detailed knowledge of the unsteady aerodynamic loads acting on the blades. In order to obtain precise insight into the aerodynamic damping of a vibrating blade assembly, a basic experiment was performed where unsteady pressure distributions were measured for subsonic and transonic flow conditions. The experiments were performed on a non-rotating, two-dimensional section of a compressor cascade in an annular test facility. The cascade consists of 20 blades (NACA3506 profile) mounted on elastic spring suspensions. In order to measure the unsteady pressure distribution, the cascade was set to tuned pitching oscillations (traveling wave modes). Each blade was driven to controlled harmonic torsional motions around midchord by a magnetic excitation system and by inductive displacement probes which measure the feedback signal of the motion. Steady and unsteady pressures were measured by steady pressure taps and piezo-electric pressure transducers, respectively. The measurement of the unsteady aerodynamic response to a shock vibrating on the suction side of the blades was enabled by a dense spacing of transducers in this region. The global aerodynamic stability is assessed by a damping coefficient evaluated from the out-of-phase parts of the unsteady moment coefficients and by the contributions from the local work coefficient, using the measured pressure data.

Sign in / Sign up

Export Citation Format

Share Document