scholarly journals Considerations on Axial Compressor Bleed for Sub-Idle Performance Models

2020 ◽  
Author(s):  
Ferran Roig Tió ◽  
Luis E. Ferrer-Vidal ◽  
Hasani Azamar Aguirre ◽  
Vassilios Pachidis
Keyword(s):  
Flow Analysis ◽  
Axial Compressor ◽  
Engine Performance ◽  
Analysis Tool ◽  
Performance Models ◽  
Performance Modelling ◽  
Start Up ◽  
Considerable Impact ◽  
Aero Engines ◽  
Analytical Approaches

Abstract The trend towards increased bypass ratio and reduced core size in civil aero-engines puts a strain on ground-start and relight capability, prompting renewed interest in sub-idle performance modelling. While a number of studies have looked at some of the broad performance modelling issues prevalent in this regime, the effects that bleed can have on sub-idle performance have not been addressed in the literature. During start-up and relight, the unknown variation in bleed flows through open handling bleed valves can have a considerable impact on the compressor’s operating line. This paper combines experimental, numerical and analytical approaches to look at the effect that sub-idle bleed flows have on predicted start-up operating lines, along with their effect on compressor characteristics. Experimental whole-engine data along with a purpose-built core-flow analysis tool are used to assess the effect of bleed model uncertainty on engine performance models. An experimental rig is used to assess the effects of reverse bleed on compressor characteristics and measurements are compared against numerical results. Several strategies for the generation of sub-idle maps including bleed effects are investigated.

scholarly journals Generating axial compressor maps to zero speed

2020 ◽  
pp. 095765092097605
Author(s):  
Luis E Ferrer-Vidal ◽  
Vassilios Pachidis ◽  
Richard J Tunstall
Keyword(s):  
Gas Turbines ◽  
Axial Compressor ◽  
Sensitivity Study ◽  
Generation Process ◽  
Performance Models ◽  
Start Up ◽  
Map Generation ◽  
Industrial Gas Turbines ◽  
Main Components ◽  
Aviation Engines

Gas turbine performance models typically rely on component maps to characterize engine component performance throughout the operational regime. For the sub-idle case, the lack of reliable rig test data or inability to run design codes far from design conditions entails that component maps have to be generated from the extrapolation of existing data at higher speeds. This undermines the accuracy of whole-engine sub-idle performance models, at times impacting engine development and certification of aviation engines and the accuracy of start-up performance prediction in industrial gas turbines. One of the main components driving this issue is the core compression system, which can present operability concerns during light-up and which also sets the combustor airflow required for ignition. This paper presents, discusses, and draws on previous approaches to describe a method enabling the creation of sub-idle compressor maps from analytical and physical grounds. The method relies on the calculation of zero-speed and torque-free lines to generate a map down to zero speed along with analytical interpolation. A method for the interpolation process is described. A sensitivity study is carried out to assess the effects that different elements of the map generation process may have on the accuracy of the resulting performance calculation. Overall, a method for the generation of accurate, consistent maps from limited geometry data is identified.

scholarly journals Axial Compressor Mean-Line Analysis: Choking Modelling and Fully-Coupled Integration in Engine Performance Simulations

2021 ◽  
Vol 6 (1) ◽  
pp. 4
Author(s):  
Ioannis Kolias ◽  
Alexios Alexiou ◽  
Nikolaos Aretakis ◽  
Konstantinos Mathioudakis
Keyword(s):  
Axial Compressor ◽  
Engine Performance ◽  
Engine Model ◽  
Multi Stage ◽  
Transient Simulations ◽  
Full Knowledge ◽  
Compressor Performance ◽  
First Time ◽  
Performance Calculation ◽  
Fully Coupled

A mean-line compressor performance calculation method is presented that covers the entire operating range, including the choked region of the map. It can be directly integrated into overall engine performance models, as it is developed in the same simulation environment. The code materializing the model can inherit the same interfaces, fluid models, and solvers, as the engine cycle model, allowing consistent, transparent, and robust simulations. In order to deal with convergence problems when the compressor operates close to or within the choked operation region, an approach to model choking conditions at blade row and overall compressor level is proposed. The choked portion of the compressor characteristics map is thus numerically established, allowing full knowledge and handling of inter-stage flow conditions. Such choking modelling capabilities are illustrated, for the first time in the open literature, for the case of multi-stage compressors. Integration capabilities of the 1D code within an overall engine model are demonstrated through steady state and transient simulations of a contemporary turbofan layout. Advantages offered by this approach are discussed, while comparison of using alternative approaches for representing compressor performance in overall engine models is discussed.

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.

scholarly journals Simplified Method for Flow Analysis and Performance Calculation Through an Axial Compressor

1984 ◽  
Author(s):  
Hubert Miton ◽  
Youssef Doumandji ◽  
Jacques Chauvin
Keyword(s):  
Equations Of Motion ◽  
Computing Time ◽  
Flow Analysis ◽  
Axial Compressor ◽  
Computation Method ◽  
Duct Flow ◽  
Axial Compressors ◽  
Flow Through ◽  
And Performance ◽  
Performance Calculation

This paper describes a fast computation method of the flow through multistage axial compressors of the industrial type. The flow is assumed to be axisymmetric between the blade rows which are represented by actuator disks. Blade row losses and turning are calculated by means of correlations. The equations of motion are linearized with respect to the log of static pressure, whose variation along the radius is usually of limited extent for the type of machines for which the method has been developed. In each computing plane (i.e. between the blade rows) two flows are combined: a basic flow with constant pressure satisfying the mass flow requirements and a perturbation flow fulfilling the radial equilibrium condition. The results of a few sample calculations are given. They show a satisfactory agreement with a classical duct flow method although the computing time is reduced by a factor five. The method has also been coupled with a surge line prediction calculation.

scholarly journals Intake Flow Analysis of a Pulsed Detonation Engine

2016 ◽  
Vol 139 (4) ◽  
Author(s):  
Joshua A. Strafaccia ◽  
Semih M. Ölçmen ◽  
John L. Hoke ◽  
Daniel E. Paxson
Keyword(s):  
Fluid Dynamic ◽  
Flow Analysis ◽  
Engine Performance ◽  
Fuel Injector ◽  
Pulse Detonation ◽  
Pulsed Detonation Engine ◽  
Pulsed Detonation ◽  
Fill Fraction ◽  
Detonation Engine ◽  
Intake Flow

Unsteady flow within the intake system of a hydrogen–air pulse detonation engine (PDE) has been analyzed using a quasi-one-dimensional (Q1D) computational fluid dynamic (CFD) code. The analysis provides insight into the unsteady nature of localized equivalence ratios and their effects on PDE performance. For this purpose, a code originally configured to model the PDE tube proper was modified to include a 6.1 m long intake with a single fuel injector located approximately 3.05 m upstream of the primary intake valve. The results show that constant fuel mass flow rate injection from the injector creates large local variations in equivalence ratio throughout the PDE within a cycle. The effect of fill fraction on the engine performance is better described with the presence of the inlet model. However, the effect of ignition delay is shown to be better predicted with a model without the inlet.

Non-Linear Aeromechanical Behaviour of Axial Core Compressor Stator Vanes

2007 ◽  
Author(s):  
Caetano Peng
Keyword(s):  
Integrated Design ◽  
Engine Performance ◽  
Forced Response ◽  
Design Tools ◽  
Virtual Testing ◽  
Aero Engine ◽  
Stator Vane ◽  
Recent Developments ◽  
Aero Engines ◽  
Engine Components

This paper highlights some engine non-linearities that can affect both performance and robustness of aero engines. It pays particular attention to non-linearities generated at the stator vane contact end joints. These non-linearities resulting from friction contact joints affect the vane modeshapes, damping and forced response. This work proposes upper and lower bound solutions based on vane end restraints non-linearities to predict conservative forced response of stator vanes. Some non-linearities such as those caused by mistuning can be beneficial to the component and system. There are also non-linearities that can be detrimental to engine performance, robustness and reliability. Moreover, it proposes and discusses the concept of temporal HCF or CCF lifing method. Recent developments in FE, CFD, mistuning, forced response and probabilistic codes can help to create more integrated design tools that incorporate time-dependent non-linearities in the lifing of aero engine components. Computations performed here demonstrated some level of component virtual testing. These analyses are important component virtual testing that will be gradually extended to whole aero engine virtual testing.

Structural Analysis of an Auxetic Casing Structure Incorporating Tip Blowing Casing Treatment Modification

2019 ◽  
Author(s):  
Tobias Schmidt ◽  
Jan Lorenz ◽  
Volker Gümmer ◽  
Andreas Hupfer
Keyword(s):  
Axial Compressor ◽  
Local Stress ◽  
Maximum Load ◽  
Element Analysis ◽  
Treatment Modification ◽  
Casing Treatment ◽  
Compressor Design ◽  
Stress Levels ◽  
Aero Engines ◽  
Casing Treatments

Abstract In axial compressor design for aero engines high system efficiency and operational stability are two main objectives. Both depend on clearance-induced losses. Previous investigations at the Institute have resulted in a passive clearance controlled compressor design using additively manufactured auxetic casing structures. The extension to an active clearance controlled device to keep an approximately constant tip gap ratio during the entire flight mission is currently investigated. Constructive on these deliverables, the implementation of tip blowing casing treatment modification in a double-walled compressor casing including an auxetic inner structure is covered in this work and studied for maximum load conditions by means of Finite Element Analysis. The idea to supplement the current auxetic casing construction with casing treatment modification emerges from the aspiration to generate further stability improvements in the high-pressure domain and the exploitation of the design freedom provided by additive manufacturing. Key issues addressed in this work by conducting parameter studies are casing treatment positioning and corresponding structural correlations depending on circumferential quantity. The evaluation section concentrates mainly on the calculated stress level associated with tip blowing casing treatments because this value is crucial for prospective fatigue predictions. In order to compare the results, the auxetic casing structure without casing treatment modification serves as reference. Promising solutions for local stress reductions are also proposed and discussed. From a structural mechanics perspective, the casing treatment modification generates very high and comparable notch stress levels at each position. Placing the casing treatments at the framework of the auxetic cells and splitting the inner casing ring results in tolerable stress levels.

Modeling Particle Deposition Effects in Aircraft Engine Compressors

2017 ◽  
Vol 139 (5) ◽  
Author(s):  
Felix Döring ◽  
Stephan Staudacher ◽  
Christian Koch ◽  
Matthias Weißschuh
Keyword(s):  
Gas Turbines ◽  
Particle Deposition ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Engine Performance ◽  
Aircraft Engine ◽  
Deposition Process ◽  
Aircraft Engines ◽  
Blade Row ◽  
Performance Deterioration

Airborne particles ingested in aircraft engines deposit on compressor blading and end walls. Aerodynamic surfaces degrade on a microscopic and macroscopic scale. Blade row, compressor, and engine performance deteriorate. Optimization of maintenance scheduling to mitigate these effects requires modeling of the deterioration process. This work provides a deterioration model on blade row level and the experimental validation of this model in a newly designed deposition test rig. When reviewing previously published work, a clear focus on deposition effects in industrial gas turbines becomes evident. The present work focuses on quantifying magnitudes and timescales of deposition effects in aircraft engines and the adaptation of the generalized Kern and Seaton deposition model for application in axial compressor blade rows. The test rig's cascade was designed to be representative of aircraft engine compressor blading. The cascade was exposed to an accelerated deposition process. Reproducible deposition patterns were identified. Results showed an asymptotic progression of blade row performance deterioration. A significant increase in total pressure loss and decrease in static pressure rise were measured. Application of the validated model using existing particle concentration and flight cycle data showed that more than 95% of the performance deterioration due to deposition occurs within the first 1000 flight cycles.

Compressor Characteristics in Gas Turbine Performance Modelling

2001 ◽  
Author(s):  
Geoff Jones ◽  
Pericles Pilidis ◽  
Barry Curnock
Keyword(s):  
Gas Turbine ◽  
Pressure Loss ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Performance Model ◽  
Flow Function ◽  
Performance Modelling ◽  
Turbine Engine ◽  
Loss Parameter ◽  
Modelling Methods

The choice of how to represent the performance of the fans and compressors of a gas turbine engine in a whole-engine performance model can be critical to the number of iterations required by the solver or indeed whether the system can be solved. This paper therefore investigates a number of compressor modelling methods and compares their relative merits. Particular attention is given to investigating the ability of the various representations to model the performance far from design point. It is noted that, for low rotational speeds and flows, matching on pressure ratio will produce problems, and that efficiency is a discontinuous function at these conditions. Thus, such traditional representations of compressors are not suitable for investigations of starting or windmilling performance. Matching on pressure ratio, Beta, the Crainic exit flow function and the true exit flow function is investigated. The independent parameters of isentropic efficiency, pressure loss, a modified pressure loss parameter, specific torque, and ideal and actual enthalpy rises are compared. The requirements of the characteristic choice are investigated, with regard to choosing matching variables and ensuring that relationships are smooth and continuous throughout the operating range of the engine.

Thermodynamics Cycle Analysis, Pressure Loss and Heat Transfer Assessment of a Recuperative System for Aero Engines

2014 ◽  
Author(s):  
A. Goulas ◽  
S. Donnerhack ◽  
M. Flouros ◽  
D. Misirlis ◽  
Z. Vlahostergios ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Fuel Consumption ◽  
Heat Exchangers ◽  
Engine Performance ◽  
Specific Fuel Consumption ◽  
Hot Gas ◽  
Research Activities ◽  
Aero Engine ◽  
Overall Efficiency ◽  
Aero Engines

Aiming in the direction of designing more efficient aero engines, various concepts have been developed in recent years, among which is the concept of an intercooled and recuperative aero engine. Particularly in the area of recuperation, MTU Aero Engines has been driving research activities in the last decade. This concept is based on the use of a system of heat exchangers mounted inside the hot-gas exhaust nozzle (recuperator). Through the operation of the system of heat exchangers, the heat from the exhaust gas, downstream the LP turbine of the jet engine is driven back to the combustion chamber. Thus, the preheated air enters the engine combustion chamber with increased enthalpy, providing improved combustion and by consequence, increased fuel economy and low-level emissions. If additionally an intercooler is placed between the compressor stages of the aero engine, the compressed air is then cooled by the intercooler thus, less compression work is required to reach the compressor target pressure. In this paper an overall assessment of the system is presented with particular focus on the recuperative system and the heat exchangers mounted into the aero engine’s exhaust nozzle. The herein presented results were based on the combined use of CFD computations, experimental measurements and thermodynamic cycle analysis. They focus on the effects of total pressure losses and heat exchanger efficiency on the aero engine performance especially the engine’s overall efficiency and the specific fuel consumption. More specifically, two different hot-gas exhaust nozzle configurations incorporating modifications in the system of heat exchangers are examined. The results show that significant improvements can be achieved in overall efficiency and specific fuel consumption hence contributing into the reduction of CO2 and NOx emissions. The design of a more sophisticated recuperation system can lead to further improvements in the aero engine efficiency in the reduction of fuel consumption. This work is part of the European funded research program LEMCOTEC (Low Emissions Core engine Technologies).

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