Unsteady Flow Simulation of Buoyancy-Driven Flows in High-Pressure Compressor Disk Cavities

2018 ◽  
Author(s):  
Atsushi Tateishi ◽  
Toshinori Watanabe ◽  
Takehiro Himeno
Keyword(s):  
High Pressure ◽  
Temperature Distribution ◽  
Unsteady Flow ◽  
Large Scale ◽  
Flow Simulation ◽  
Detached Eddy Simulation ◽  
Hot Air ◽  
High Pressure Compressor ◽  
Compressor Disk ◽  
Pressure Compressor

This paper focuses on the buoyancy-induced unsteady flow phenomenon inside high-pressure compressor disk cavities. In order to understand the flow structure in a realistic configuration, a 10-stage core compressor of the NASA/GE Energy Efficient Engine is adopted as a computational target. The numerical flow simulation is conducted on a full annulus model, where the temperature distribution on the wall is modeled based on the core test results. The time-averaged flow fields are obtained by detached eddy simulation (DES) and two-dimensional axisymmetric Reynolds-averaged Navier-Stokes (RANS) simulation, and the difference is discussed in detail. The DES result showed large-scale, vortical structures with significant radial velocity fluctuations especially in the rear part of the compressor. These fluctuations create radial arm-like structure in the temperature distribution in the cavity, and greatly enhance the mixing between the bore coolant and hot air near the cavity wall. In addition, it is observed that the hot air discharged from the cavities creates a large cell at bore region, which extends across several rear stages. Although the present study successfully illustrates the entire structure of unsteady flow in heated compressor disk cavities including full stages, a more detailed validation will be needed to further confirm the applicability of DES for the targeted flow.

The Use of CFD for the Thermal Analysis of a High Pressure Compressor Rotor

2004 ◽  
Author(s):  
U. W. Ruedel ◽  
J. R. Turner
Keyword(s):  
Thermal Analysis ◽  
High Pressure ◽  
Temperature Distribution ◽  
Flow Field ◽  
Transient Temperature ◽  
Temperature Profiles ◽  
Cfd Analysis ◽  
Compressor Rotor ◽  
High Pressure Compressor ◽  
Pressure Compressor

The prediction of fatigue life of components inside aircraft engines depends on the reliable numerical modelling of the temperature distribution during a mission cycle as this gives rise to life limiting thermal stresses. The transient temperature distribution is usually measured during an engine test and is then used to validate the numerical model, which in turn produces the basis for calculating the thermal stress levels. This paper describes the thermal analysis of a High Pressure Compressor Rotor (HPCR) and how the use of a 3-D Computational Fluid Dynamic (CFD) analysis improved the quantitative agreement between the measured and the predicted temperature profiles. The highly complex three-dimensional flow field within the compressor rotor was modelled by exploiting symmetry conditions and using a standard k-ε turbulence model. Results of the tangential, axial and radial velocity components as well as locations of peaks in turbulence kinetic energy were predicted to help identify the flow field inside the forward cavity of the rotor. Two ways of predicting internal re-circulating rates to the rim area are proposed. Finally, plots of predicted metal temperature profiles before and after the CFD-analysis are presented.

scholarly journals DEVELOPMENT OF THE NUMERICAL METHODOLOGY FOR THE ESTIMATION OF RESONANT PROCESSES AT GASTURBINE ENGINE FLOW DUCT

2020 ◽  
pp. 12-19
Author(s):  
Nikolay Shuvaev ◽  
◽  
Aleksandr Siner ◽  
Ruslan Kolegov ◽  
◽  
...  
Keyword(s):  
High Pressure ◽  
Unsteady Flow ◽  
Aircraft Engine ◽  
Operating Conditions ◽  
Design Stage ◽  
High Pressure Compressor ◽  
Wide Range ◽  
Resonant Processes ◽  
Flow Duct ◽  
Pressure Compressor

Ensuring safety of flights is the most important task that is being solved in the process of designing an aircraft engine and aircraft. The most complex are the physical processes occurring inside the aircraft engine, especially in its gas generator: combustion chamber, high-pressure compressor and high-pressure turbine. The unsteady flow of gas in the flow duct of the aircraft engine is very complex, it is difficult to model, because the flow is characterized by a wide range of time and space scales. Unsteady flow in a high-pressure compressor can cause surge and breakdown of the compressor and the entire engine as a whole. Along with the detachment flows causing the surge, in the flow duct there can be resonant phenomena associated with the propagation of powerful sound waves along the flow duct of the engine, which, when a direct and reflected wave is imposed, create a very powerful standing wave that affects the structure. With a certain combination of conditions, the coincidence of the natural frequencies of the oscillations of the air volume and the solid body, such resonant processes in the flow duct of the gas turbine engine can lead to serious breakdowns, such as breakage of rotor blades and guide vanes, destruction of the aeroengine framework and other. The main difficulty is that it is problematic to identify such processes at the design and debugging stage, since there are no suitable mathematical models, and for experimental verification it is required to withstand the specific operating conditions of the node that are not known in advance. This work is devoted to the creation of a calculation technique that will allow in the future to diagnose resonance phenomena at the design stage and thereby significantly reduce the costs for the design, testing and manufacture of an aircraft engine. The proposed technique is based on the nonstationary Navier-Stokes equations for a compressible gas.

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.

A Machine Learning Based Approach of Performance Estimation for High-Pressure Compressor Airfoils

2018 ◽  
Author(s):  
Jonas Marx ◽  
Stefan Gantner ◽  
Jörn Städing ◽  
Jens Friedrichs
Keyword(s):  
Machine Learning ◽  
High Pressure ◽  
Machine Learning Algorithms ◽  
Performance Estimation ◽  
Predictive Maintenance ◽  
Support Vector ◽  
K Nearest Neighbor ◽  
Operating Characteristics ◽  
High Pressure Compressor ◽  
Pressure Compressor

In recent years, the demands of Maintenance, Repair and Overhaul (MRO) customers to provide resource-efficient after market services have grown increasingly. One way to meet these requirements is by making use of predictive maintenance methods. These are ideas that involve the derivation of workscoping guidance by assessing and processing previously unused or undocumented service data. In this context a novel approach on predictive maintenance is presented in form of a performance-based classification method for high pressure compressor (HPC) airfoils. The procedure features machine learning algorithms that establish a relation between the airfoil geometry and the associated aerodynamic behavior and is hereby able to divide individual operating characteristics into a finite number of distinct aero-classes. By this means the introduced method not only provides a fast and simple way to assess piece part performance through geometrical data, but also facilitates the consideration of stage matching (axial as well as circumferential) in a simplified manner. It thus serves as prerequisite for an improved customary HPC performance workscope as well as for an automated optimization process for compressor buildup with used or repaired material that would be applicable in an MRO environment. The methods of machine learning that are used in the present work enable the formation of distinct groups of similar aero-performance by unsupervised (step 1) and supervised learning (step 2). The application of the overall classification procedure is shown exemplary on an artificially generated dataset based on real characteristics of a front and a rear rotor of a 10-stage axial compressor that contains both geometry as well as aerodynamic information. In step 1 of the investigation only the aerodynamic quantities in terms of multivariate functional data are used in order to benchmark different clustering algorithms and generate a foundation for a geometry-based aero-classification. Corresponding classifiers are created in step 2 by means of both, the k Nearest Neighbor and the linear Support Vector Machine algorithms. The methods’ fidelities are brought to the test with the attempt to recover the aero-based similarity classes solely by using normalized and reduced geometry data. This results in high classification probabilities of up to 96 % which is proven by using stratified k-fold cross-validation.

Elimination of unstable free vibrations of the rotor of a centrifugal high-pressure compressor

1988 ◽  
Vol 24 (7) ◽  
pp. 356-360
Author(s):  
V. B. Shnepp ◽  
A. M. Galeev ◽  
G. S. Batkis ◽  
V. M. Polyakov
Keyword(s):  
High Pressure ◽  
Free Vibrations ◽  
High Pressure Compressor ◽  
Pressure Compressor

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.

scholarly journals Acquisition of F-100(3) High Pressure Compressor Entrance Profiles

1982 ◽  
Author(s):  
D. C. Rabe ◽  
W. W. Copenhaver ◽  
M. S. Perry
Keyword(s):  
High Pressure ◽  
Total Pressure ◽  
System Component ◽  
Sensing Element ◽  
Flow Angle ◽  
Automatic Data ◽  
Turbine Engine ◽  
High Pressure Compressor ◽  
Air Force Base ◽  
Pressure Compressor

A transportable automatic data acquisition system to obtain high pressure compressor entrance profiles in an F-100 Series 3 gas turbine engine is described. The system was developed, assembled, and tested at Wright-Patterson Air Force Base and transported to a remote location for implementation in a sea level engine test. Acquisition of data was controlled through a Hewlett Packard Model 9825T desktop calculator, preprogrammed to display airflow data in engineering units during the test. Entrance profiles of total and static pressure, temperature, and flow angle for two axial locations are presented. A wedge probe sensing element was positioned at 12 radial locations by remote traversing mechanisms to obtain these profiles. For a total pressure range of 18 to 46 psia (0.13 to 0.32 MPa), acquisition uncertainties in static and total pressure were reduced to below ± percent of measured values by optimizing data system component uncertainties.

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