Effects of Compressor Tip Injection on Aircraft Engine Performance and Stability

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
Wolfgang Horn ◽  
Klaus-Jürgen Schmidt ◽  
Stephan Staudacher
Keyword(s):  
High Pressure ◽  
Steady State ◽  
Mass Flow ◽  
Engine Performance ◽  
Operational Characteristic ◽  
Control Logic ◽  
High Pressure Compressor ◽  
Tip Injection ◽  
Stability Enhancement ◽  
Pressure Compressor

This analytical study discusses the system aspects of active stability enhancement using mass flow injection in front of the rotor blade tip of a high pressure compressor. Tip injection is modeled as a recirculating bleed in a performance simulation of a commercial turbofan engine. A map correction procedure accounts for the changes in compressor characteristics caused by injection. The correction factors are derived from stage stacking calculations, which include a simple correlation for stability enhancement. The operational characteristic of the actively controlled engine is simulated in steady and transient states. The basic steady-state effect consists of a local change in mass flow and a local increase in gas temperature. This alters the component matching in the engine. The mechanism can be described by the compressor-to-turbine flow ratio and the injection temperature ratio. Both effects reduce the cycle efficiency resulting in an increased turbine temperature and fuel consumption at constant thrust. The negative performance impact becomes negligible if compressor recirculation is only employed at the transient part power and if valves remain closed at the steady-state operation. Detailed calculations show that engine handling requirements and temperature limits will still be met. Tip injection increases the high pressure compressor stability margin substantially during critical maneuvers. The proposed concept in combination with an adequate control logic offers promising benefits at transient operation, leading to an improvement potential for the overall engine performance.

Effects of Compressor Tip Injection on Aircraft Engine Performance and Stability

2007 ◽  
Author(s):  
Wolfgang Horn ◽  
Klaus-Ju¨rgen Schmidt ◽  
Stephan Staudacher
Keyword(s):  
High Pressure ◽  
Steady State ◽  
Mass Flow ◽  
Engine Performance ◽  
Operational Characteristic ◽  
Control Logic ◽  
High Pressure Compressor ◽  
Tip Injection ◽  
Stability Enhancement ◽  
Pressure Compressor

This analytical study discusses the system aspects of active stability enhancement using mass flow injection in front of the rotor blade tip of a high pressure compressor. Tip injection is modeled as a recirculating bleed in a performance simulation of a commercial turbofan engine. A map correction procedure accounts for the changes in compressor characteristics caused by injection. The correction factors are derived from stage stacking calculations which include a simple correlation for stability enhancement. The operational characteristic of the actively controlled engine is simulated in steady and transient states. The basic steady-state effect consists of a local change in mass flow and a local increase in gas temperature. This alters the component matching in the engine. The mechanism can be described by the compressor-to-turbine flow ratio and the injection temperature ratio. Both effects reduce the cycle efficiency resulting in increased turbine temperature and fuel consumption at constant thrust. The negative performance impact becomes negligible if compressor recirculation is only employed at transient part power and if valves remain closed at steady-state operation. Detailed calculations show that engine handling requirements and temperature limits will still be met. Tip injection increases the high pressure compressor stability margin substantially during critical maneuvers. The proposed concept in combination with an adequate control logic offers promising benefits at transient operation, leading to an improvement potential for overall engine performance.

High Pressure Compressor Stabilization by Controlled Pulsed Injection

2012 ◽  
Author(s):  
Sven-Jürgen Hiller ◽  
Jakob Hermann ◽  
Robert Widhopf-Fenk
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Stability Control ◽  
Stable Operation ◽  
Surge Margin ◽  
High Pressure Compressor ◽  
Stability Control System ◽  
Pulsed Injection ◽  
Tip Injection ◽  
Pressure Compressor

Steady tip injection with re-circulated air as a reliable method for high pressure compressor stabilization has the disadvantage to reduce the efficiency due to the usage of already compressed air. Therefore, the minimization of recirculation air and/or a large extension of the stable operation range is mandatory. The paper describes an alternative approach by using an active compressor stability control system with modulating valves. The system was tested at a high pressure compressor. It recognized evolving stall cells reliable. The achieved available stable throttling range of the compressor (in terms of mass flow) was more than doubled by active controlled injection compared to steady tip injection at 90% aerodynamic speed. The other way round, depending on the tip flow structure and the aerodynamic speed injection mass flow saving up to 50% at a similar surge margin compared to steady injection was achieved.

scholarly journals Analisis Pengaruh High Pressure Compressor Rotor Clearance Terhadap Exhaus Gas Temperature Margin pada CFM56-7

2021 ◽  
Vol 6 (2) ◽  
pp. 50-55
Author(s):  
Wildan Sofary Darga ◽  
Edy K. Alimin ◽  
Endah Yuniarti
Keyword(s):  
High Pressure ◽  
Engine Performance ◽  
Exhaust Gas ◽  
Gas Temperature ◽  
Compressor Rotor ◽  
High Pressure Compressor ◽  
The Difference ◽  
Gas Turbine Exhaust ◽  
Temperature Margin ◽  
Pressure Compressor

Exhaust Gas Temperatue is an parameter where the hot gases’s temperature leave the gas turbine. Exhaust gas temperature margin is the difference between highest temperature at take off phase with redline on indicator (???????????? ???????????????????????? °????=???????????? ????????????????????????????−???????????? ???????????????? ????????????). EGTM is one of any factor to determine engine performance. A good perfomance of an engine when it has a big margin (EGTM), during operation of an engine the EGTM could decrease untill 0 (zero). So many factors could affect EGTM deteroration there are: distress hardware such as airfoil erosion, leak of an airseals, and increase of clearance between tip balde and shroud. Increase of clearance happens in high pressure compressor rotor clearance. In CFM56-7 have 9 stage(s) of high pressure compressor and each stage give the EGT Loses. The calculation of EGT Effect/Losses is actual celarance – minimum clearance x 1000 x EGT Effect °C, where actual clearance define by the substraction of outside diameter’s rotor with inside diameter’s shroud, minimum clearance define in the manual, 1000 is adjustment from mils/microinch to inch, and EGT Effect is temperature that define in the manual. The analysist had done with 6 (six) engine serial number and proceed by corelation that shown linkage between clearance and EGT Effect, the corelation is strong shown the result of corelation (r) is 0.994275999 or nearest 1.

A Robust Design Methodology for High-Pressure Compressor Throughflow Optimization

2003 ◽  
Author(s):  
N. Lecerf ◽  
D. Jeannel ◽  
A. Laude
Keyword(s):  
High Pressure ◽  
Standard Deviation ◽  
Robust Design ◽  
Design Methodology ◽  
Engine Performance ◽  
Design Parameters ◽  
Deterministic Optimization ◽  
High Pressure Compressor ◽  
Noise Factors ◽  
Pressure Compressor

Reducing costs and development times are two of the main challenges for aircraft engines manufacturers. Analysis shows that the main troubles encountered during the industrialization phase are due to choices made during the first steps, such as the preliminary design of the compressor throughflow (flowpath and velocity triangles). Therefore, constraints and needs from the later phases have to be taken into account as early as possible. A deterministic optimization method for automated compressor throughflow design has been developed to achieve these objectives, improving efficiency and surge margin while modifying the design parameters. Nevertheless, variability between the theoretical geometry and the actual one may occur because of the manufacturing process or the damages encountered during the engine life cycle. Depending on their magnitude, these differences can affect the engine performance. To consider these random phenomena from the design step, the deterministic optimization is coupled with a probabilistic approach, based on a robust design methodology which aims at guarantee the engine performance despite geometrical variability. This article deals with geometrical robustness. It presents a robust design methodology and introduces a capability function used to optimize the outputs of a compressor model while minimizing their standard deviation. The model has two kinds of inputs: the design factors, which are known by both designer and manufacturer, and the noise factors, that are just known by their mean value and their standard deviation. As robust design requires a large number of calculations, it is interesting to work with an approximated physical model such as a response surface, generated through the computation of a suitable design of experiments. This method has been successfully applied to the design of a Snecma Moteurs high-pressure compressor.

Accuracy Assessment of Steady and Unsteady Multistage High Pressure Compressor Simulations

2019 ◽  
Author(s):  
Sebastian Lück ◽  
Stefan Kuntzagk ◽  
Guido Doebbener ◽  
Andreas Kellersmann ◽  
Christoph Bode ◽  
...  
Keyword(s):  
High Pressure ◽  
Steady State ◽  
Accuracy Assessment ◽  
Flow Characteristics ◽  
Time Resolved ◽  
Engine Model ◽  
High Pressure Compressor ◽  
Stage Performance ◽  
Steady State Simulation ◽  
Pressure Compressor

Abstract In this paper a comparison of the performance parameters and flow characteristics of a mature commercial high bypass engine’s 9-stages high pressure compressor (HPC) with steady-state mixing-plane (RANS) and unsteady RANS (URANS) CFD is carried out. The investigation is based on a numerical model of the CFM56-5C engine’s HPC which is installed on the Airbus A340-300 aircraft. The compressor under investigation features the so called 3D blading which is the first of two performance improvement packages available. An experimental engine of this type equipped with extensive additional instrumentation is in use by Lufthansa Technik to deliver detailed experimental data of the gas path. Experimental results have been discussed in previous works of the authors. In order to provide long-term forecasts of an engine’s state the aforementioned CFD model has been included into a multilevel engine model. To be able to evaluate the significance and applicability of the CFD results into such model, it is of great interest to which extent and level of detail it can deliver accurate performance predictions. From the comparison of both steady-state and unsteady simulation results it is found that overall compressor performance only differs negligibly while stage performance can differ significantly. It is depicted that among the stator vanes of the front stages local supersonic flow and flow separation can occur. These are not captured by the steady-state simulation to the same degree a time resolved simulation does. In rear stages differences fade and unsteady methods tend to predict better stage performance which may be due to favourable effects of rotor-stator interaction.

Effects of tip injection on the performance of a multi-stage high-pressure compressor

2011 ◽  
Vol 2 (1-4) ◽  
pp. 99-110 ◽  
Author(s):  
M. Kern ◽  
W. Horn ◽  
S.-J. Hiller ◽  
S. Staudacher
Keyword(s):  
High Pressure ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Tip Injection ◽  
Pressure Compressor

Review on the aerodynamics of intermediate compressor duct

2020 ◽  
Vol 14 (4) ◽  
pp. 7446-7468
Author(s):  
Manish Sharma ◽  
Beena D. Baloni
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Flow Analysis ◽  
Outer Wall ◽  
Optimization Techniques ◽  
Optimum Pressure ◽  
Research Areas ◽  
Static Pressure Distribution ◽  
High Pressure Compressor ◽  
Pressure Compressor

In a turbofan engine, the air is brought from the low to the high-pressure compressor through an intermediate compressor duct. Weight and design space limitations impel to its design as an S-shaped. Despite it, the intermediate duct has to guide the flow carefully to the high-pressure compressor without disturbances and flow separations hence, flow analysis within the duct has been attractive to the researchers ever since its inception. Consequently, a number of researchers and experimentalists from the aerospace industry could not keep themselves away from this research. Further demand for increasing by-pass ratio will change the shape and weight of the duct that uplift encourages them to continue research in this field. Innumerable studies related to S-shaped duct have proven that its performance depends on many factors like curvature, upstream compressor’s vortices, swirl, insertion of struts, geometrical aspects, Mach number and many more. The application of flow control devices, wall shape optimization techniques, and integrated concepts lead a better system performance and shorten the duct length.  This review paper is an endeavor to encapsulate all the above aspects and finally, it can be concluded that the intermediate duct is a key component to keep the overall weight and specific fuel consumption low. The shape and curvature of the duct significantly affect the pressure distortion. The wall static pressure distribution along the inner wall significantly higher than that of the outer wall. Duct pressure loss enhances with the aggressive design of duct, incursion of struts, thick inlet boundary layer and higher swirl at the inlet. Thus, one should focus on research areas for better aerodynamic effects of the above parameters which give duct design with optimum pressure loss and non-uniformity within the duct.

scholarly journals Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions

2014 ◽  
Author(s):  
Alain Batailly ◽  
Mathias Legrand ◽  
Antoine Millecamps ◽  
Sèbastien Cochon ◽  
François Garcin
Keyword(s):  
High Pressure ◽  
Forced Response ◽  
Compressor Blade ◽  
Design Stage ◽  
Blade Design ◽  
Early Integration ◽  
Nonlinear Structural ◽  
High Pressure Compressor ◽  
Structural Interactions ◽  
Pressure Compressor

Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural blade/casing contacts at the design stage of a high-pressure compressor blade. The blade of interest underwent structural divergence after blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the blade stacking law while maintaining aerodynamic performance of the original blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.

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