Study on a Quasi-Two-Dimensional Fan Model for Variant Bypass-Ratio Condition

2018 ◽  
Vol 0 (0) ◽  
Author(s):  
Dawei Fu ◽  
Yong Wang ◽  
Haibo Zhang ◽  
Qiangang Zheng
Keyword(s):  
Pressure Ratio ◽  
Flow Characteristics ◽  
Axial Flow ◽  
Ratio Error ◽  
Compensation Factor ◽  
Radial Profiles ◽  
Ratio Condition ◽  
Iteration Program ◽  
Nominal Performance ◽  
Bypass Ratio

Abstract In order to describe how the variant BPR affects the flow performance, a qusi-2D fan model was proposed. In the model, the stage cumulative method was applied to compute axial flow and the simplified radial equilibrium equation was applied to compute radial flow, then a dual-level iteration program was proposed for the nominal performance calculation. To compute the non-nominal flow characteristics on non-nominal BPR, a triple-level iteration program with the compensation factor was proposed to obtain the radial profiles. On that basis, Blade angles were designed for this model on the design point. With the inflow deflection factor designed as 0.8 and the loss factor along blade as 0.95, the pressure ratio error is only 0.21 % and the isentropic efficiency error is 0.06 %. Simulation showed that the variant-BPR profiles and the overall characteristics of flow at fan exit or for bypass & core ducts could be well computed with much less time-consuming.

Surge-Induced Structural Loads in Gas Turbines

1980 ◽  
Vol 102 (1) ◽  
pp. 162-168 ◽  
Author(s):  
R. S. Mazzawy
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Turbine Engine ◽  
Compression System ◽  
Engine Design ◽  
Steady State Levels ◽  
Flow Compression ◽  
Engine Components ◽  
Bypass Ratio

The axial flow compression system of a modern gas turbine engine normally delivers a large quantity of airflow at relatively high velocity. The sudden stoppage (and reversal) of this flow when an engine surges can result in structural loads in excess of steady state levels. These loads can be quite complex due to inherent asymmetry in the surge event. The increasing requirements for lighter weight engine structures, coupled with the higher pressure ratio cycles required for minimizing fuel consumption, make the accurate prediction of these loads an important part of the engine design process. This paper is aimed toward explaining the fluid mechanics of the surge phenomenon and its impact on engine structures. It offers relatively simple models for estimating surge-induced loads on various engine components. The basis for these models is an empirical correlation of surge-induced inlet overpressure based on engine pressure ratio and bypass ratio. An approximate estimate of the post-surge axial pressure distribution can be derived from this correlation by assuming that surge initiation occurs in the rear of the compression system.

Aerodynamic Design of a Highly Loaded Axial Flow Fan Rotor Using a Novel One-Dimensional Design Method With its Numerical Simulation

2016 ◽  
Author(s):  
Ali Shahsavari ◽  
Mahdi Nili-Ahamadabadi
Keyword(s):  
Numerical Simulation ◽  
Design Method ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Numerical Results ◽  
Meridional Plane ◽  
Performance Curve ◽  
One Dimensional ◽  
Bypass Ratio

This paper presents a novel one-dimensional design method based on the radial equilibrium theory and constant span-wise diffusion factor to redesign of NASA rotor 67 just aerodynamically with a higher pressure ratio at the same design point. A one-dimensional design code is developed to obtain the meridional plane and blade to blade geometry of rotor to reach the three-dimensional view of rotor blades. To verify the redesigned rotor, its flow numerical simulation is carried out to compute its performance curve. The experimental performance curve of NASA rotor 67 is used for validation of the numerical results. Structured mesh with finer grids near walls is used to capture flow field and boundary layer effects. RANS equations are solved by finite volume method for rotating zones and stationary zones. The numerical results of the new rotor show about 9% increase in its pressure ratio at both design and off design mass flow rate. The new rotor has a higher outlet velocity through its upper span improving bypass ratio of a turbofan engine. To prove the new fan ability of producing more bypass ratio, a thermodynamic analysis is conducted. The results of this analysis show 13% increase in bypass ratio and 5.7% decline in specific fuel consumption in comparison to NASA rotor 67.

A High-Loaded Supersonic Axial Compressor Aerodynamic Design Principle

2012 ◽  
Author(s):  
Yingjiao Hu ◽  
Songtao Wang
Keyword(s):  
Numerical Simulation ◽  
Design Principle ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Axial Compressor ◽  
Axial Flow ◽  
Internal Flow ◽  
Aerodynamic Design ◽  
Simulation Results

Reviewed the historical development of the supersonic axial flow compressor, and gave an outlook for its future developments and research orientations. According to the internal flow characteristics of the conventional supersonic axial flow compressors, put forward a high load of supersonic axial compressor aerodynamic design principle. A preliminary design verification of the principle has been carried. The 3D viscous numerical simulation results show that, under the tip tangential speed 360m / s, has achieved a stage pressure ratio 2.3 with efficiency 86.5%. In addition, considering the rotor under impulse condition can get the maximum rotor total pressure ratio with high efficiency, a design principle has also been put forward to solve the high entrance Mach number problem of the downstream stator. But the numerical simulation results show that the multi-shock structure does not have any advantages to reduce the stator losses.

The Flow Characteristics of a Piston-Cylinder Assembly with an Off-Centre, Open Port

1980 ◽  
Vol 194 (1) ◽  
pp. 291-299 ◽  
Author(s):  
A. P. Morse ◽  
J. H. Whitelaw ◽  
M. Yianneskis
Keyword(s):  
Cylinder Head ◽  
Laser Doppler Anemometry ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Axial Flow ◽  
Circumferential Velocity ◽  
Piston Cylinder ◽  
Radial Profiles ◽  
Intake Stroke ◽  
Open Port

Laser-Doppler anemometry has been used to measure the mean and r.m.s. values of the velocity components of the flow in a reciprocating assembly containing an off-centre, open port in the cylinder head and a flat piston. The port was inclined at an angle of 60° to the head and gave rise to a three-dimensional flow with no significant compression. The rotational speed was 200 rev/min, ensuring fully turbulent flow. The measurements were restricted to the axial and circumferential velocity components and were made on three radial planes in order to describe the flow in sufficient detail. The results are presented in the form of radial profiles and vector diagrams in both the axial and cross-stream planes. The main flow pattern on the intake stroke is of the same general shape as that previously obtained with a symmetric port, a large vortex occupying the space bounded by the indrawn jet with a much smaller vortex in the corner between the cylinder head and wall. These vortices are however no longer symmetrical about the cylinder axis, but vary in both size and intensity with circumferential position. In the cross-stream plane, the primary vortex is characterized by the presence of a stress-induced swirling motion on either side of the plane of symmetry with velocities significantly greater than the circumferential velocity of the jet. At mid-stroke, these vortices rotate counter to the jet, but at 144° ATDC, the direction of rotation reverses over part of the length of the vortex and is in the same sense as the jet motion. On the exhaust stroke, the flow patterns are relatively simple with quasi-parallel axial flow and little swirl until the port is approached, when the outgoing jet accelerates and begins to rotate in a direction opposite to its motion on the intake stroke.

scholarly journals Axial Flow Compressor Stability Enhancement: Circumferential T-Shape Grooves Performance Investigation

Aerospace ◽  
2021 ◽  
Vol 8 (1) ◽  
pp. 12
Author(s):  
Marco Porro ◽  
Richard Jefferson-Loveday ◽  
Ernesto Benini
Keyword(s):  
Vortex Breakdown ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Suction Side ◽  
Treatment Technique ◽  
Tip Leakage Flow ◽  
Stall Inception ◽  
Stall Margin ◽  
Casing Treatment ◽  
Stall Margin Improvement

This work focuses its attention on possibilities to enhance the stability of an axial compressor using a casing treatment technique. Circumferential grooves machined into the case are considered and their performances evaluated using three-dimensional steady state computational simulations. The effects of rectangular and new T-shape grooves on NASA Rotor 37 performances are investigated, resolving in detail the flow field near the blade tip in order to understand the stall inception delay mechanism produced by the casing treatment. First, a validation of the computational model was carried out analysing a smooth wall case without grooves. The comparisons of the total pressure ratio, total temperature ratio and adiabatic efficiency profiles with experimental data highlighted the accuracy and validity of the model. Then, the results for a rectangular groove chosen as the baseline case demonstrated that the groove interacts with the tip leakage flow, weakening the vortex breakdown and reducing the separation at the blade suction side. These effects delay stall inception, improving compressor stability. New T-shape grooves were designed keeping the volume as a constant parameter and their performances were evaluated in terms of stall margin improvement and efficiency variation. All the configurations showed a common efficiency loss near the peak condition and some of them revealed a stall margin improvement with respect to the baseline. Due to their reduced depth, these new configurations are interesting because they enable the use of a thinner light-weight compressor case as is desirable in aerospace applications.

Experimental research on the performance of the forward variable area bypass injector for variable cycle engines

2020 ◽  
Vol 0 (0) ◽  
Author(s):  
Binglong Zhang ◽  
He Liu ◽  
Jianhua Zhou ◽  
Hui Liu
Keyword(s):  
Pressure Ratio ◽  
Effective Means ◽  
Experimental System ◽  
Area Ratio ◽  
Back Pressure ◽  
Mode Transition ◽  
Bypass Valve ◽  
Injection Coefficient ◽  
Bypass Ratio ◽  
Specific Working

AbstractThe forward variable area bypass injector (FVABI) is a key component of double bypass variable cycle engine (VCE) to achieve mode transition and bypass ratio adjustment. In this paper, an experimental system for FVABI was constructed based on the analysis of relevant experimental theories, and then the experiments on FVABI were carried out for a specific working state in double bypass mode of VCE and for the comparison working states with different area ratios and different back pressure ratios. The results showed that the FVABI designed in this paper meets the requirements of VCE at this working state. The analysis of the influence of area ratio and back pressure ratio on the injection coefficient showed that the first bypass valve and back pressure were effective means to control the mass flow of FVABI.

Tip Leakage Flow Characteristics Downstream of an Axial Flow Fan

1996 ◽  
Author(s):  
P. Puddu
Keyword(s):  
Vortex Flow ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Axial Flow ◽  
Tip Leakage Flow ◽  
Single Wire ◽  
Vortex System ◽  
Tip Leakage ◽  
Stress Components ◽  
Computer Controlled

The three-dimensional viscous flow characteristics and the complex vortex system downstream of the rotor of an industrial exial fan have been determined by an experimental investigation using hot-wire anemometer. Single-wire slanted and straight type probes have been rotated about the probe axis using a computer controlled stepper motor. Measurements have been taken at four planes behind the blade trailing edge. The results show the characteristics of the relative flow as velocity components, secondary flow and kinetic energy defect. Turbulence intensity and Reynolds stress components in the leakage vortex area are also presented. The evolution of the leakage vortex flow during the decay process has also been evaluated in terms of dimension, position and intensity.

Ultra High Bypass Ratio Engine Sizing and Cycle Selection Study for a Subsonic Commercial Aircraft in the N+2 Timeframe

2011 ◽  
Author(s):  
Brian K. Kestner ◽  
Jeff S. Schutte ◽  
Jonathan C. Gladin ◽  
Dimitri N. Mavris
Keyword(s):  
Pressure Ratio ◽  
Important Variable ◽  
Direct Drive ◽  
Commercial Aircraft ◽  
Low Pressure Turbine ◽  
Fuel Burn ◽  
Technology Transition ◽  
Fundamental Research ◽  
Bypass Ratio ◽  
Engine Cycle

This paper presents an engine sizing and cycle selection study of ultra high bypass ratio engines applied to a subsonic commercial aircraft in the N+2 (2020) timeframe. NASA has created the Environmentally Responsible Aviation (ERA) project to serve as a technology transition bridge between fundamental research (TRL 1–4) and potential users (TRL 7). Specifically, ERA is focused on subsonic transport technologies that could reach TRL 6 by 2020 and are capable of integration into an advanced vehicle concept that simultaneously meets the ERA project metrics for noise, emissions, and fuel burn. An important variable in exploring the trade space is the selection of the optimal engine cycle for use on the advanced aircraft. In this paper, two specific ultra high bypass engine cycle options will be explored: advanced direct drive and geared turbofan. The advanced direct drive turbofan is an improved version of conventional turbofans. In terms of both bypass ratio and overall pressure ratio, the advanced direct turbofan benefits from improvements in aerodynamic design of its components, as well as material stress and temperature properties. By putting a gear between the fan and the low pressure turbine, a geared turbo fan allows both components to operate at optimal speeds, thus further improving overall cycle efficiency relative to a conventional turbofan. In this study, sensitivity of cycle design with level of technology will be explored, in terms of both cycle parameters (such as specific thrust consumption (TSFC) and bypass ratio) and aircraft mission parameters (such as fuel burn and noise). To demonstrate this sensitivity, engines will be sized for optimal performance on a 300 passenger class aircraft for a 2010 level technology tube and wing airframe, a N+2 level technology tube and wing air-frame, and finally on a N+2 level technology blended wing body airframe with and without boundary layer ingestion (BLI) engines.

Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

2017 ◽  
Author(s):  
Justin (Jongsik) Oh
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Single Stage ◽  
Design Flow ◽  
Design Parameters ◽  
Original Design ◽  
Circular Arc ◽  
Surge Margin ◽  
Flow Compressor

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

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