Exergetic and Energetic Response Surfaces for Small Turbojet Engine

2011 ◽  
Vol 110-116 ◽  
pp. 1054-1058
Author(s):  
Onder Turan ◽  
T. Hikmet Karakoc
Keyword(s):  
Mach Number ◽  
Centrifugal Compressor ◽  
Exergy Analysis ◽  
Reverse Flow ◽  
Three Dimensional ◽  
Axial Flow ◽  
Engine Performance ◽  
Response Surfaces ◽  
Operating Parameters ◽  
Turbojet Engine

Exergy analysis permits meaningful efficiencies to be evaluated for a system or process, and the sources, causes and locations of thermodynamic losses to be determined. This study presents exergetic modeling of a small turbojet engine via exergetic response surfaces. Turbojet engine consists of an inlet, a centrifugal compressor, reverse flow combustion chamber, axial-flow turbine and exhaust nozzle. The flight Mach number and altitude are examined on the exergetic efficiencies of total engine performance. The results of analysis are given as three dimensional exergetic response surface plots related to these operating parameters.

Exergetic Destruction Effects of Operating Conditions on the Turbojet Engine Components

2011 ◽  
Vol 110-116 ◽  
pp. 2390-2394
Author(s):  
T. Hikmet Karakoc ◽  
Onder Turan
Keyword(s):  
Mach Number ◽  
Combustion Chamber ◽  
Reverse Flow ◽  
Three Dimensional ◽  
Axial Flow ◽  
Operating Conditions ◽  
Exergy Destruction ◽  
Theoretical Limit ◽  
Turbojet Engine ◽  
Engine Components

The minimization of exergy destruction brings the design as closely as permissible to the theoretical limit. This study presents exergy destruction analysis of a turbojet engine for different flight Mach number and altitudes. Turbojet engine being considered consists of an inlet, a centrifugal compressor, reverse flow combustion chamber, axial-flow turbine and exhaust nozzle. The flight Mach number and altitude are examined on the exergetic destructions of compressor, combustion chamber, turbine and exhaust nozzle. The results of component-based destruction analysis are given as three dimensional exergetic-destruction response surface plots related to altitude and flight Mach number.

Compressor Performance 2D Modelling at Reverse Flow Conditions

2010 ◽  
Author(s):  
L. Gallar ◽  
I. Tzagarakis ◽  
V. Pachidis ◽  
R. Singh
Keyword(s):  
Experimental Data ◽  
Reverse Flow ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Engine Performance ◽  
Two Dimensional ◽  
Flow Conditions ◽  
Compressor Surge ◽  
Compressor Performance ◽  
Shaft Failure

After a shaft failure the compression system of a gas turbine is likely to surge due to the heavy vibrations induced on the engine after the breakage. Unlike at any other conditions of operation, compressor surge during a shaft over-speed event is regarded as desirable as it limits the air flow across the engine and hence the power available to accelerate the free turbine. It is for this reason that the proper prediction of the engine performance during a shaft over-speed event claims for an accurate modelling of the compressor operation at reverse flow conditions. The present study investigates the ability of the existent two dimensional algorithms to simulate the compressor performance in backflow conditions. Results for a three stage axial compressor at reverse flow were produced and compared against stage by stage experimental data published by Gamache. The research shows that due to the strong radial fluxes present over the blades, two dimensional approaches are inadequate to provide satisfactory results. Three dimensional effects and inaccuracies are accounted for by the introduction of a correction parameter that is a measure of the pressure loss across the blades. Such parameter is tailored for rotors and stators and enables the satisfactory agreement between calculations and experiments in a stage by stage basis. The paper concludes with the comparison of the numerical results with the experimental data supplied by Day on a four stage axial compressor.

Effects of Inlet Flow Field Conditions on the Performance of Centrifugal Compressor Diffusers: Part 2 — Straight-Channel Diffuser

1998 ◽  
Author(s):  
Sabri Deniz ◽  
Edward M. Greitzer ◽  
Nicholas A. Cumpsty
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Centrifugal Compressor ◽  
Axial Flow ◽  
Rotating Stall ◽  
Pressure Recovery ◽  
Straight Channel ◽  
Inlet Flow ◽  
Flow Angle ◽  
Operating Range

This is Part 2 of an examination of influence of inlet flow conditions on the performance and operating range of centrifugal compressor vaned diffusers. The paper describes tests of straight-channel type diffuser, sometimes called a wedge-vane diffuser, and compares the results with those from the discrete-passage diffusers described in Part 1. Effects of diffuser inlet Mach number, flow angle, blockage, and axial flow non-uniformity on diffuser pressure recovery and operating range are addressed. The straight-channel diffuser investigated has 30 vanes and was designed for the same aerodynamic duty as the discrete-passage diffuser described in Part 1. The ranges of the overall pressure recovery coefficients were 0.65–0.78 for the straight-channel diffuser and 0.60–0.70 for the discrete-passage diffuser; the pressure recovery of the straight-channel diffuser was roughly 10% higher than that of the discrete-passage diffuser. Both types of the diffusers showed similar behavior regarding the dependence on diffuser inlet flow angle and the insensitivity of the performance to inlet flow field axial distortion and Mach number. The operating range of the straight-channel diffuser, as for the discrete-passage diffusers was limited by the onset of rotating stall at a fixed momentum-averaged flow angle into the diffuser, which was for the straight-channel diffuser, αcrit = 70° ±0.5°. The background, nomenclature and description of the facility and method are all given in Part 1.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Vol 126 (4) ◽  
pp. 473-481 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and Computational Fluid Dynamics (CFD). This report presents the measured results of the high transonic flow at the impeller inlet using Laser Doppler Velocimeter (LDV) and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution, and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460 m/s and the relative Mach number reaches about 1.6. Using a LDV, about 500 m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and CFD. This report presents the measured results of the high transonic flow at the impeller inlet using LDV and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460m/s and the relative Mach number reaches about 1.6. Using an LDV, about 500m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Radial Ultra-Micro Wave Rotors (UµWR): Design and Simulation

2006 ◽  
Author(s):  
Florin Iancu ◽  
Janusz Piechna ◽  
Norbert Mu¨ller
Keyword(s):  
Reverse Flow ◽  
Three Dimensional ◽  
Axial Flow ◽  
Wave Rotor ◽  
Compression Process ◽  
Radial Wave ◽  
Wave Rotors ◽  
Commercial Code ◽  
Through Flow ◽  
Computational Fluid Dynamics Cfd

It has been shown that the wave rotor technology has the potential of improving the performance of gas turbine cycles. Moreover the radial wave rotor is an additional innovation for this technology. Unlike the commercialized axial-flow wave rotor (Comprex®), a radial one has the benefit of using centrifugal forces to improve the compression process or flow scavenging. The geometry of the rotor is much simpler and is ideal for microfabrication, which is relying mainly on two-dimensional processes to create three-dimensional features. This paper is presenting several radial ultra-micro wave rotors (UμWR) configurations and numerical analysis of these rotors. In a radial placement, the wave rotor has four possible configurations: two - general configuration, through-flow and reverse-flow, and each of these could have the low pressure air port positioned at inside or outside of the rotor. Results have been obtained using FLUENT, a Computational Fluid Dynamics (CFD) commercial code. The vast information about the unsteady processes occurring during simulation is visualized.

A Fast Gas Turbine Engine Performance Analysis Tool Able to Capture Three Dimensional Effects

2007 ◽  
Author(s):  
Ioannis Templalexis ◽  
Pericles Pilidis ◽  
Vassilios Pachidis ◽  
Petros Kotsiopoulos
Keyword(s):  
Three Dimensional ◽  
Axial Flow ◽  
Engine Performance ◽  
Flow Simulation ◽  
Simulation Method ◽  
Flow Profile ◽  
Flow Distortion ◽  
Performance Simulation ◽  
Turbine Engine ◽  
The Impact

Given the current level of computational resources that are readily available, three dimensional (3-D) gas turbine engine performance simulation remains extremely time consuming. The current paper presents a synthesis of existing flow simulation methods coupled together in the form of a new software package. The software is able to assess the impact of a 3-D flow profile at the intake inlet on engine performance, demanding relatively low computational resources. More precisely four flow simulation techniques are employed, represented respectively by four individual stand alone software sub-modules. 3-D Vortex Lattice Method (VLM) is used to simulate the intake flow. Subsequently the intake outlet 3-D flow profile is decomposed into a radial and a circumferential component. For the compressor performance simulation, that receives those components as inlet boundary conditions, a two dimensional (2-D) Streamline Curvature (SLC) simulation method coupled with an extended parallel compressor model is used. SLC addresses the impact of the radial flow distortion, whereas the extended parallel compressor model examines the impact of circumferential flow distortion on engine performance. The results of the above analysis are stored into an intake-compressor performance characteristic map, which is then fed into a zero dimensional (0-D) performance simulation tool in order to evaluate the overall impact of the intake inlet distorted flow on engine performance. The paper is divided into two major sections. The first one presents the individual flow simulation techniques, together with the corresponding software modules. A short summary of each method is given first and then the software module is described, followed by brief comments on the validation results that have been already published. The section in concluded by the description of the synthesized software. The second major section deals with the application of the synthesized simulation method on a turbojet engine. A generic turbojet engine has been chosen mounted behind a generic intake, given the lack of relevant experimental results. The engine has a four stage axial flow compressor driven by a single stage axial flow turbine, followed by a converging nozzle. 3-D total pressure profiles were imposed at the intake inlet and several comparative graphs of engine’s performance parameters between “clean” and distorted inlet flow conditions are given. The paper is concluded with a discussion on software’s abilities and weaknesses as well as on its potential future expansion.

On-Line Compressor Cascade Washing for Gas Turbine Performance Investigation

2011 ◽  
Author(s):  
Uyioghosa Igie ◽  
Pericles Pilidis ◽  
Dimitrios Fouflias ◽  
Ken Ramsden ◽  
Paul Lambart
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Engine Performance ◽  
Compressor Cascade ◽  
Design Point ◽  
Washing Process ◽  
On Line ◽  
Flow Compressor

On-line compressor washing for industrial gas turbine application is a promising method of mitigating the effects of compressor fouling degradation; however there are still few studies from actual engine experience that are inconclusive. In some cases the authors attribute this uncertainty as a result of other existing forms of degradation. The experimental approach applied here is one of the first of its kind, employing on-line washing on a compressor cascade and then relating the characteristics to a three-dimensional axial flow compressor. The overall performance of a 226MW engine model for the different cases of a clean, fouled and washed engine is obtained based on the changing compressor behavior. Investigating the effects of fouling on the clean engine exposed to blade roughness of 102μm caused 8.7% reduction in power at design point. This is equivalent, typically to 12 months degradation in fouling conditions. Decreases in mass flow, compressor efficiency, pressure ratio and unattainable design point speed are also observed. An optimistic recovery of 50% of the lost power is obtained after washing which lasts up to 10mins. Similarly, a recovery of all the key parameters is achieved. The study provides an insight into compressor cascade blade washing, which facilitates a reliable estimation of compressor overall efficiency penalties based on well established assumptions. Adopting Howell’s theory as well as constant polytropic efficiency, a general understanding of turbomachinery would judge that 50% of lost power recovered is likely to be the high end of what is achievable for the existing high pressure wash. This investigation highlights the obvious benefits of power recovery with on-line washing and the potential to maintain optimum engine performance with frequent washes. Clearly, the greatest benefits accrue when the washing process is initiated immediately following overhaul.

Loss Mechanisms and Flow Control for Improved Efficiency of a Centrifugal Compressor at High Inlet Prewhirl

2015 ◽  
Author(s):  
Zheng Xinqian ◽  
Huang Qiangqiang ◽  
Liu Anxiong
Keyword(s):  
Flow Control ◽  
Mass Flow ◽  
Centrifugal Compressor ◽  
Reverse Flow ◽  
Three Dimensional ◽  
Leading Edge ◽  
Flow Region ◽  
Percentage Points ◽  
Stable Conditions ◽  
Loss Mechanisms

Variable inlet prewhirl is an effective way to suppress compressor instability. Compressors usually employ a high degree of positive inlet prewhirl to shift the surge line in the performance map to a lower mass flow region. However, the efficiency of a compressor at high inlet prewhirl is far lower than that at zero or low prewhirl. This paper investigates the performances of a centrifugal compressor with different prewhirl, discusses the mechanisms thought to be responsible for the production of extra loss induced by high inlet prewhirl and develops flow control methods to improve efficiency at high inlet prewhirl. The approach combines steady three-dimensional Reynolds average Navier-Stockes (RANS) simulations with theoretical analysis and modeling. In order to make the study universal to various applications with inlet prewhirl, the inlet prewhirl was modeled by modifying the velocity condition at the inlet boundary. Simulation results show that the peak efficiency at high inlet prewhirl is reduced compared to that at zero prewhirl by over 7.6 percentage points. The extra loss is produced upstream and downstream of the impeller. Severe flow separation was found near the inlet hub which reduces efficiency by 2.3 percentage points. High inlet prewhirl works like a centrifuge gathering low-kinetic-energy fluid to hub, inducing the separation. A dimensionless parameter C is defined to measure the centrifugal component of flow. As for the extra loss produced downstream of the impeller, the flow mismatch of impeller and diffuser at high prewhirl causes a violent backflow near the diffuser vanes’ leading edges. An analytical model is built to predict diffuser choking mass flow which proves that the diffuser flow operates outside of stable conditions. Based on the two loss mechanisms, hub curve and diffuser stager angle were modified and adjusted for seeking higher efficiency at high prewhirl. The efficiency improvement of a modification of the hub is 1.1 percentage points and that of the combined optimization is 2.4 percentage points. During optimizing, constant distribution of inlet prewhirl was found to induce reverse flow at the leading edge of the blade root, which turned out being uncorrelated with blade angle. By revealing loss mechanisms and proposing flow control ideas, this paper lays a theoretical basis for overcoming the efficiency drop induced by high inlet prewhirl and for developing compressors with high inlet prewhirl.

Reverse Flow Turbine-Like Operation of an Axial Flow Compressor

2012 ◽  
Author(s):  
A. Gill ◽  
T. W. von Backström ◽  
T. M. Harms
Keyword(s):  
Reverse Flow ◽  
Flow Direction ◽  
Pressure Ratio ◽  
Harmonic Approximation ◽  
Three Dimensional ◽  
Axial Flow ◽  
Time Dependent ◽  
Navier Stokes ◽  
Axial Flow Compressor ◽  
Flow Compressor

This article describes an experimental investigation of the flow structures occurring in an axial flow compressor during second quadrant operation for reversed rotor rotation in the incompressible flow regime. In second quadrant operation, the flow direction is reversed, but the pressure is lower at the compressor inlet than at the outlet. The compressor thus acts as an axial flow turbine. A three stage axial flow compressor, with a mass flow rate of 2.7 kg/s and a pressure ratio of 1.022 was investigated. The design rotor tip Mach number is 0.2. Three operational points within the second quadrant were investigated, at flow coefficients of −0.482, −0.553 and −0.843. A five hole conical probe and a 50 micron diameter inclined hot film anemometer were used in this investigation. Radial traverses downstream of rotor rows and a time-dependent area traverse downstream of the first stage stator were performed. Three-dimensional time-dependent numerical Navier-Stokes solutions using the non-linear harmonic approximation for single blade passages in each blade row for each of the cases are compared with experimental work. The compressor has already been show to be capable of attaining relatively high turbine efficiency (76%) when operating in this mode. Examination of the flow field shows that little to no flow separation occurs on the rotor or stator blades. The wakes of all blades are found to be thin and sharp, and the area between wakes is large and approximately uniform. Numerical results agree relatively well with experimental results.

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