scholarly journals Testing and Performance Verification of a High Bypass Ratio Turbofan Rotor in an Internal Flow Component Test Facility

2007 ◽  
Author(s):  
Dale E. Van Zante ◽  
Gary G. Podboy ◽  
Christopher J. Miller ◽  
Scott A. Thorp
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Flow Field ◽  
Pressure Rise ◽  
Internal Flow ◽  
Lessons Learned ◽  
Scale Model ◽  
Test Facility ◽  
Inlet Flow ◽  
Bypass Ratio

A 1/5 scale model rotor representative of a current technology, high bypass ratio, turbofan engine was installed and tested in the W8 single-stage, high-speed, compressor test facility at NASA Glenn Research Center (GRC). The same fan rotor was tested previously in the GRC 9×15 Low Speed Wind Tunnel as a fan module consisting of the rotor and outlet guide vanes mounted in a flight-like nacelle. The W8 test verified that the aerodynamic performance and detailed flow field of the rotor as installed in W8 were representative of the wind tunnel fan module installation. Modifications to W8 were necessary to ensure that this internal flow facility would have a flow field at the test package that is representative of flow conditions in the wind tunnel installation. Inlet flow conditioning was designed and installed in W8 to lower the fan face turbulence intensity to less than 1.0% in order to better match the wind tunnel operating environment. Also, inlet bleed was added to thin the casing boundary layer to be more representative of a flight nacelle boundary layer. On the 100% speed operating line the fan pressure rise and mass flow rate agreed with the wind tunnel data to within 1%. Detailed hot film surveys of the inlet flow, inlet boundary layer and fan exit flow were compared to results from the wind tunnel. The effect of inlet casing boundary layer thickness on fan performance was quantified. Challenges and ‘lessons learned’ from testing this high flow, low static pressure rise fan in an internal flow facility are discussed.

Experimental Investigation of the Effect of the Probe Support Tail Structure on the Compressor Cascade Flow Field

2019 ◽  
Vol 36 (1) ◽  
pp. 9-18
Author(s):  
Honghui Xiang ◽  
Ning Ge ◽  
Jie Gao ◽  
Rongfei Yang ◽  
Minjie Hou
Keyword(s):  
Experimental Investigation ◽  
Flow Field ◽  
Flow Condition ◽  
Test Facility ◽  
Axial Distance ◽  
Compressor Cascade ◽  
Inlet Flow ◽  
Disturbance Intensity ◽  
Cascade Flow ◽  
Tail Structure

Abstract Aiming at resolving the problem of measuring probe blockage effect in the performance experiments of high loaded axial flow compressors, an experimental investigation of the probe support disturbance effect on the compressor cascade flow field was conducted on a transonic plane cascade test facility. The influence characteristics of the probe support tail structure on the cascade downstream flow field under different operation conditions were revealed through the detailed analysis of the test data. The results show that the aerodynamic coupling effect between the upstream probe support wake and the downstream cascade flow field is very intense. Some factors, i. e. inlet Mach number, probe support tail structure, circumferential installing position of probe, and axial distance from the probe support trailing edge to the downstream cascade, are found to have the most impact on the probe disturbance intensity. Under high speed inlet flow condition, changing probe support tail structure can’t inhibit probe support disturbance intensity effectively. Whereas under low speed inlet flow condition, compared with the cylindrical probe, the elliptic probe can inhibit probe support wake loss and reduce disturbance effects on the downstream cascade flow field.

The Influence of Fan Root Flow on the Aerodynamics of a Low-Pressure Compressor Transition Duct

2019 ◽  
Author(s):  
A. D. Walker ◽  
I. Mariah ◽  
D. Tsakmakidou ◽  
H. Vadhvana ◽  
C. Hall
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Inlet Flow ◽  
Flat Profile ◽  
Bypass Ratio

Abstract To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.

Effects of Inlet Flow Field Conditions on the Stall Onset of Centrifugal Compressor Vaned Diffusers

2003 ◽  
Author(s):  
Sabri Deniz
Keyword(s):  
Flow Field ◽  
Axial Flow ◽  
Rotating Stall ◽  
Field Conditions ◽  
Test Facility ◽  
Straight Channel ◽  
Inlet Flow ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
Operating Range

This paper considers the performance and operating range of vaned diffusers for use in high performance centrifugal compressors. An experimental and numerical investigation is performed to determine the effects of inlet flow field conditions on pressure recovery and stall onset of different type vaned diffusers, such as discrete-passage and straight-channel diffusers. Diffuser inlet flow conditions examined include Mach number, flow angle, blockage, and axial flow non-uniformity. The investigation was carried out in a specially built test facility, designed to provide a controlled inlet flow field to the test diffusers. Unsteady pressure measurements showed the operating range of a compressor stage was limited by the onset of rotating stall, triggered by the loss of stability in the vaned diffuser, independent of the impeller operating point. For both diffusers investigated, loss of flow stability in the diffuser occurred at a critical value of the momentum-averaged flow angle into the diffuser. To provide additional information on diffuser flow development and to complement previous experimental work performed on straight-channel type diffuser, a computational investigation has been undertaken and important results are presented.

Surface Flow Patterns as Visualised by Dust Deposits on the Blades of a Fan

1963 ◽  
Vol 67 (633) ◽  
pp. 589-594 ◽  
Author(s):  
E. T. Hignett ◽  
M. M. Gibson
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Leading Edge ◽  
Surface Flow ◽  
Scale Model ◽  
Flow Visualisation ◽  
Pressure Gradients ◽  
Suction Surface ◽  
Layer Flow ◽  
Dust Deposits

Investigations by one of the authors in connection with the design of a fan for a blower type of wind tunnel showed that regular and repeatable dust patterns occurred on the blades of a one-quarter scale model fan of 18 inches diameter. Dust was deposited on the fan blades along the leading-edge and on the suction surface over an area thought to be the turbulent region of the boundary layer. The introduction of isolated protuberances on the dust free area of a blade gave rise to turbulence wedges in which dust was also deposited and this was interpreted as confirmation of the coincidence of the dust deposits with regions of turbulent boundary-layer flow. These deposits showed the existence of a considerable extent of laminar flow on the suction surface of each blade close to the root, a region where high lift coefficients would be expected with associated adverse pressure gradients. Two-dimensional wind tunnel experiments were made to confirm the interpretation of the observed dust patterns by comparison with the smoke filament and volatile liquid methods of flow visualisation and these are reported in Reference 2.

scholarly journals Effect of Inlet Geometry on Fan Performance and Flow Field in a Half-Ducted Propeller Fan

2012 ◽  
Vol 2012 ◽  
pp. 1-9 ◽  
Author(s):  
Pin Liu ◽  
Norimasa Shiomi ◽  
Yoichi Kinoue ◽  
Ying-zi Jin ◽  
Toshiaki Setoguchi
Keyword(s):  
Flow Field ◽  
Flow Measurement ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Internal Flow ◽  
Fan Performance ◽  
Ducted Propeller ◽  
Inlet Geometry ◽  
Blade Tip ◽  
The One

In order to clarify the effect of rotor inlet geometry of half-ducted propeller fan on performance and velocity fields at rotor outlet, the experimental investigation was carried out using a hotwire anemometer. Three types of inlet geometry were tested. The first type is the one that the rotor blade tip is fully covered by a casing. The second is that the front one-third part of blade tip is opened and the rest is covered. The third is that the front two-thirds are opened and the rest is covered. Fan test and internal flow measurement at rotor outlet were conducted about three types of inlet geometry. At the internal flow measurement, a single slant hotwire probe was used and a periodical multisampling technique was adopted to obtain the three-dimensional velocity distributions. From the results of fan test, the pressure-rise characteristic drops at high flowrate region and the stall point shifts to high flowrate region, when the opened area of blade tip increases. From the results of velocity distributions at rotor outlet, the region with high axial velocity moves to radial inwards, the circumferential velocity near blade tip becomes high, and the flow field turns to radial outward, when the opened area increases.

Wind effects on rough-walled and smooth-walled large cooling towers

2016 ◽  
Vol 20 (6) ◽  
pp. 843-864 ◽  
Author(s):  
XX Cheng ◽  
L Zhao ◽  
YJ Ge ◽  
R Dong ◽  
C Demartino
Keyword(s):  
Boundary Layer ◽  
Surface Roughness ◽  
Wind Tunnel ◽  
Flow Field ◽  
Atmospheric Boundary Layer ◽  
Model Test ◽  
Full Scale ◽  
Cooling Towers ◽  
Test Results ◽  
Wind Effects

Adding vertical ribs is recognized as a useful practice for reducing wind effects on cooling towers. However, ribs are rarely used on cooling towers in China since Chinese Codes are insufficient to support the design of rough-walled cooling towers, and an “understanding” hampers the use of ribs, which thinks that increased surface roughness has limited effects on the maximum internal forces that control the structural design. To this end, wind tunnel model tests in both uniform flow field with negligible free-stream turbulence and atmospheric boundary layer (ABL) turbulent flow field are carried out in this article to meticulously study and quantify the surface roughness effects on both static and dynamic wind loads for the purpose of improving Chinese Codes first. Subsequently, a further step is taken to obtain wind effects on a full-scale large cooling tower at a high Re, which are employed to validate the results obtained in the wind tunnel. Finally, the veracity of the model test results is discussed by investigating the Reynolds number (Re) effects on them. It has been proved that the model test results for atmospheric boundary layer flow field are all obtained in the range of Re-independence and the conclusions drawn from model tests and full-scale measurements basically agree, so most model test results presented in this article can be directly applied to the full-scale condition without corrections.

scholarly journals The Model V84.3 Shot Tests: Compressor Flow Field Measurements and Evaluation

1994 ◽  
Author(s):  
M. Janssen ◽  
R. Mönig ◽  
J. Seume ◽  
H. Hönen ◽  
R. Lösch-Schloms ◽  
...  
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Static Pressure ◽  
Pressure Rise ◽  
Full Scale ◽  
Test Facility ◽  
Design Calculation ◽  
Flow Measurements ◽  
Flow Angle ◽  
Probe Measurements

Detailed experimental investigations were carried out at the Siemens test-facility in Berlin to validate and develop further the compressor design of the Model V84.3 gas turbine and to generate a comprehensive data base for the verification of the flow calculation programs. The test facility enables Siemens to confirm the design with regard to performance and reliability in the full scale machine under full load and off-design condition. Various measuring techniques well established in the laboratory were applied to the full scale compressor to examine the flow field. Along with rather conventional 5-hole probes for measuring the flow field in the core region, miniaturized 3-hole probes were developed at the Turbomachinery Laboratory of the Technical University of Aachen, tested and finally used for the measurements of endwall boundary layer profiles and their development throughout the compressor. In addition to the probe measurements, wall static-pressure measurements, as well as probed vane measurements, were carried out. The paper briefly describes the test facility, the compressor under investigation, and the instrumentation for the flow measurements. A comparison of the 3-hole and 5-hole probe measurements is presented. The experimental results are compared with calculated results taken from a two-dimensional off-design calculation program with standard loss models. By means of the measured static-pressure rise at the casing wall and the total pressure distributions downstream of the rotor rows, a modification of the loss modeling was performed. The calculated flow field is compared to the results of the 3-hole and 5-hole probe measurements in terms of radial distributions for flow angle. Mach number and total pressure.

scholarly journals Test Facility for High-Speed Probe Calibration

2019 ◽  
Vol 213 ◽  
pp. 02033
Author(s):  
Tomáš Jelínek ◽  
Erik Flídr ◽  
Martin Němec ◽  
Jan Šimák
Keyword(s):  
Wind Tunnel ◽  
Flow Field ◽  
Total Pressure ◽  
High Speed ◽  
Test Facility ◽  
Nozzle Outlet ◽  
Wide Range ◽  
Mach Numbers ◽  
Probe Calibration ◽  
Subsonic Nozzle

A new test facility was built up as a part of a closed-loop transonic wind tunnel in VZLU´s High-speed Aerodynamics Department. The wind tunnel is driven by a twelve stage radial compressor and Mach and Reynolds numbers can be changed by the compressor speed and by the total pressure in the wind tunnel loop by a set of vacuum pumps, respectively. The facility consists of an axisymmetric subsonic nozzle with an exit diameter de = 100 mm. The subsonic nozzle is designed for regimes up to M = 1 at the nozzle outlet. At the nozzle inlet there is a set of a honeycomb and screens to ensure the flow stream laminar at the outlet of the nozzle. The subsonic nozzle can be supplemented with a transonic slotted nozzle or a supersonic rigid nozzle for transonic and supersonic outlet Mach numbers. The probe is fixed in a probe manipulator situated downstream of the nozzle and it ensures a set of two perpendicular angles in a wide range (±90°). The outlet flow field was measured through in several axial distances downstream the subsonic nozzle outlet. The total pressure and static pressure was measured in the centreline and the total pressure distribution in the vertical and horizontal plane was measured as well. Total pressure fluctuations in the nozzle centreline were detected by a FRAP probe. From the initial flow measurement in a wide range of Mach numbers the best location for probe calibration was chosen. The flow field was found to be suitable for probe calibration.

Current Forces on Prelude Hull and Water Intake Riser by CFD: Full Scale Model Using Two of the World’s Largest Supercomputers

2021 ◽  
Author(s):  
Erwan Auburtin ◽  
Jang Kim ◽  
Hyunchul Jang ◽  
Lawrence Lai ◽  
Jason McConochie ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Wind Tunnel ◽  
Water Intake ◽  
Full Scale ◽  
Scale Model ◽  
Accurate Estimation ◽  
Correct Prediction ◽  
Wind Tunnel Tests ◽  
Numerical Predictions

Abstract The Prelude Floating Liquefied Natural Gas (FLNG) facility is moored with an internal turret allowing it to perform offloading operations of liquefied natural and petroleum gas products. It does so in either a Free Weathervaning (FW) mode, i.e. by allowing the unit to rotate according to environmental loads, or in a Thruster-Assisted (TA) mode, i.e. by using the stern thrusters to maintain a fixed heading deemed preferable for the entire operation, or a particular phase. An accurate estimation of the various environment effects, in terms of forces on the FLNG and LNG carrier, is critical to ensure a correct prediction of its heading or the required thruster forces, depending on the selected operating mode. The predominant loads driving the weathervaning behavior are wind and current loads. These loads have been estimated from wind tunnel tests during the engineering phase. Since the Prelude FLNG has been installed on-site, field measurements have provided an opportunity for comparison and shown some differences with the numerical predictions based on the estimated loads, prompting a need for verification of current loads by an independent method. For the Prelude FLNG application, current loads play an important role due to facility size and significant tidal currents. It has been shown in some previous studies that wind tunnel tests for a model of under-water geometry may underestimate current loads compared to those on a full-scale vessel. There is a boundary layer along the wind tunnel floor in wind tunnel tests, while the current profile is relatively uniform over the hull draft in the real ocean condition. Moreover experimental tests present some additional drawbacks: they are performed at a reduced scale (1:225), the Reynolds number is lower than full-scale even with a large wind tunnel speed, and it is difficult to model the long (150m full-scale) Water Intake Risers (WIR) extending below the hull bottom. In order to investigate these effects, state-of-the-art full-scale CFD simulations were performed for the Prelude hull and WIR. The test program included different current speeds and directions, and several sensitivity studies: Reynolds number effect between model- and full-scales, effect of current speed profile (comparing uniform and boundary layer profiles at model scale), effect of FLNG rotation in yaw, impact of unsteady current, and presence of marine growth. Extreme dimensions of Prelude FLNG and requirements for accuracy of this study called for the CFD calculations to be performed on the High Performance Computing (HPC) clusters - Stampede2 and Frontera - at the Texas Advanced Computing Center (TACC), which are both amongst the world’s largest supercomputers. This paper describes the assumptions and challenges of the CFD study and discusses the results of the main program and various sensitivities. The main conclusions and lessons learnt are also discussed.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

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