scholarly journals Turbine Map Extension - Theoretical Considerations and Practical Advice

2020 ◽  
Vol 4 ◽  
pp. 176-189
Author(s):  
Kurzke Joachim
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Flow Region ◽  
Aircraft Engine ◽  
Special Role ◽  
Power Simulation ◽  
Business Jet ◽  
Flow Limit ◽  
Geometrical Data

Physically sound compressor and turbine maps are the key to accurate aircraft engine performance simulations. Usually, maps only cover the speed range between idle and full power. Simulation of starting, windmilling and re-light requires maps with sub-idle speeds as well as pressure ratios less than unity. Engineers outside industry, universities and research facilities may not have access to the measured rig data or the geometrical data needed for CFD calculations. Whilst research has been made into low speed behavior of turbines, little has been published and no advice is available on how to extrapolate maps. Incompressible theory helps with the extrapolation down to zero flow as in this region the Mach numbers are low. The zero-mass flow limit plays a special role; its shape follows from turbine velocity triangle analysis. Another helpful correlation is how mass flow at a pressure ratio of unity changes with speed. The consideration of velocity triangles together with the enthalpy-entropy diagram leads to the conclusion that in these circumstances flow increases linearly with speed. In the incompressible flow region, a linear relationship exists between torque/flow and flow. The slope is independent of speed and can be found from the speed lines for which data are available. This knowledge helps in extending turbine maps into the regions where pressure ratio is less than unity. The application of the map extension method is demonstrated with an example of a three-stage low pressure turbine designed for a business jet engine.

Turbocharger-Design Effects on Gasoline-Engine Performance

2005 ◽  
Vol 127 (3) ◽  
pp. 525-530 ◽  
Author(s):  
Theodosios Korakianitis ◽  
T. Sadoi
Keyword(s):  
Steady Flow ◽  
Mass Flow ◽  
Gasoline Engine ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Final Choice ◽  
Design Point ◽  
Theoretical Considerations ◽  
Flow Experiments

Specification of a turbocharger for a given engine involves matching the turbocharger performance characteristics with those of the piston engine. Theoretical considerations of matching turbocharger pressure ratio and mass flow with engine mass flow and power permits designers to approach a series of potential turbochargers suitable for the engine. Ultimately, the final choice among several candidate turbochargers is made by tests. In this paper two types of steady-flow experiments are used to match three different turbochargers to an automotive turbocharged-intercooled gasoline engine. The first set of tests measures the steady-flow performance of the compressors and turbines of the three turbochargers. The second set of tests measures the steady-flow design-point and off-design-point engine performance with each turbocharger. The test results show the design-point and off-design-point performance of the overall thermodynamic cycle, and this is used to identify which turbocharger is suitable for different types of engine duties.

Effects of Number of Circumferential Casing Grooves on Stall Flow Characteristics of a Transonic Axial Compressor

2013 ◽  
Vol 284-287 ◽  
pp. 727-732
Author(s):  
Jin Hyuk Kim ◽  
Kwang Yong Kim ◽  
Kyung Hun Cha
Keyword(s):  
Mass Flow ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Flow Characteristics ◽  
Axial Compressor ◽  
Flow Region ◽  
High Mass ◽  
Convergence Criteria ◽  
Transonic Axial Compressor

This work investigates the effects of circumferential casing grooves on stall flow characteristics of a transonic axial compressor. Numerical analysis is conducted by solving three-dimensional steady Reynolds-averaged Navier-Stokes equations with the shear stress transport turbulence model. The results of flow analysis for an axial compressor with smooth casing are validated in comparison with experimental data for the pressure ratio and adiabatic efficiency. The numerical stall inception point is identified from the last converged point by convergence criteria, and the stall margin is predicted numerically. The peak adiabatic efficiency point is also obtained by reducing the normalized mass flow in the high mass flow region. In order to explore the influence of number of the circumferential casing grooves on the performance of the compressor, the stall margins and peak adiabatic efficiencies are evaluated compared to the case smooth casing. The stability of the axial compressor with circumferential casing grooves is found to be sensitively influenced by the number of grooves.

scholarly journals Ideal Power Output — An Alternative Aircraft Engine Performance Comparator

1997 ◽  
Author(s):  
Marvin F. Schmidt ◽  
Christopher M. Norden ◽  
Jeffrey M. Stricker
Keyword(s):  
Gas Turbine ◽  
Power Output ◽  
State Of The Art ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Aircraft Engine ◽  
The State ◽  
Power Cycle ◽  
Shaft Power ◽  
Bypass Ratio

The gas turbine is applied in four basic configurations; the turbojet, the turbofan, the turboprop and the turboshaft. Comparisons of the performance of these various configurations is difficult since they convert the energy to different forms, i.e. thrust or shaft power. Cycle variables which do not necessarily constitute advancements in the state-of-the-art such as bypass ratio and fan pressure ratio can have a profound effect on thrust and shaft power. Differences in flight speed and altitude capability further confound the comparisons. What is required is a comparison methodology that removes all of these variables and yet puts all the various types of engines on an equitable basis. This paper will provide such a comparison tool. All turbomachinery, regardless of configuration, can be compared with this method.

Turbocharger-Design Effects on Gasoline-Engine Performance

1997 ◽  
Author(s):  
T. Korakianitis ◽  
T. Sadoi
Keyword(s):  
Steady Flow ◽  
Mass Flow ◽  
Gasoline Engine ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Final Choice ◽  
Design Point ◽  
Theoretical Considerations ◽  
Flow Experiments

Specification of a turbocharger for a given engine involves matching the turbocharger performance characteristics with those of the piston engine. Theoretical considerations of matching turbocharger pressure ratio and mass flow with engine mass flow and power permits designers to approach a series of potential turbochargers suitable for the engine. Ultimately, the final choice among several candidate turbochargers is made by tests. In this paper two types of steady-flow experiments are used to match three different turbochargers to an automotive turbocharged-intercooled gasoline engine. The first set of tests measures the steady-flow performance of the compressors and turbines of the three turbochargers. The second set of tests measures the steady-flow design-point and off-design-point engine performance with each turbocharger. The test results show the design-point and off-design-point performance of the over-all thermodynamic cycle, and this is used to identify which turbocharger is suitable for different types of engine duties.

scholarly journals Stage reduced counter-rotating axial compressor for jet engine

2021 ◽  
Vol 12 ◽  
pp. 3
Author(s):  
F. Ferdaus ◽  
Nitish Kumar ◽  
G. Sakthivel ◽  
N. Raghukiran
Keyword(s):  
Control System ◽  
Mass Flow ◽  
Working Condition ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Aircraft Engine ◽  
Maximum Pressure ◽  
Jet Engine ◽  
Main Factors ◽  
The Stability

Variation in the states of system, mass flow and pressure are some of the disturbances which are experienced by the compressors in the jet engine under working condition. One of the main factors that influence the efficiency of a jet engine is the pressure ratio. In order to achieve the required pressure ratio, we should have relatively a greater number of stages in the compressor that leads to an increase in the weight of the engine. The stator and rotor are the essential parts of an aircraft's axial compressor. CFD is used in order to evaluate the pressure ratio. In this paper, we are going to analyze a three-stage compressor instead of an actual six-stage compressor. The mass flow rate inside the control system can be used to maintain the stability of the system. Compressor weight and pressure ratio at each stage can be reduced if we have a clockwise and anti-clockwise rotating rotor. With the use of a universal gear system, the two clockwise rotors and one anti-clockwise rotor were analyzed. The main outlook of this work is to show the maximum pressure ratio of the compressor at the outlet with our desired configurations. In conclusion, it was shown that the weight of the aircraft engine can be effectively reduced.

Effects of Ice Formation on the Flowfield of an Aircraft Engine Inlet

2008 ◽  
Author(s):  
Ray R. Taghavi ◽  
Wonjin Jin
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Total Pressure ◽  
Free Stream ◽  
Flow Region ◽  
Aircraft Engine ◽  
Ice Formation ◽  
Ice Accretion ◽  
Mach Numbers

The effects of typical rime and glaze ice on the performance of the M2129 S-duct inlet are computationally investigated using the steady-state RANS solution. The glaze ice accretion produces a substantial degradation of the inlet performance due to its obstructive shape to the in-flow, while the effect of the rime ice is not significant. Compared to the clean inlet, the secondary flow region at the engine face of the duct inlet is increased by 600 percent for the glaze iced inlet. Total pressure recoveries at the engine face for the rime and glaze ice case are 98.8 and 95.8 percent, respectively. Also, the glaze ice causes 26 percent reduction in the mass flow rate at the engine face. In addition, the adverse effects on the performance of the inlet are enhanced by an increase in freestream Mach numbers due to the stronger and more extensive shock formations in the inlet flow. With increasing free stream Mach numbers from M∞ = 0.13 to 0.85, total pressure recovery decreases from 0.985 to 0.62 with the glaze ice accretion. And the level of the mass flow rate with the glaze ice accretion is 76 percent of that in ice-free condition at M∞ = 0.13; however, it decreases to 68 percent at M∞ = 0.475.

Fluidics in Aircraft Engine Controls

1981 ◽  
Vol 103 (4) ◽  
pp. 324-330 ◽  
Author(s):  
G. E. Davies
Keyword(s):  
Pressure Ratio ◽  
Engine Performance ◽  
Aircraft Engine ◽  
Appropriate Technology ◽  
Engine Fuel ◽  
Engine Control System ◽  
Fuel Flow ◽  
Basic Engine ◽  
Wide Range ◽  
Compressor Inlet

Fluidics is a particularly appropriate technology for engine controls as it is capable of measuring pressure ratio, one of the basic engine performance parameters, as a fundamental quantity. This results in improved accuracy of measurement and obviates the need of conventional systems to utilize nonoptimum, but easier to measure, control parameters. A wide range of aircraft engine controls has been developed covering controls for compressor inlet guide vanes, compressor bleed valves, engine fuel flow including engine instrumentation. Total fluidic unit deliveries exceed 4200 and the civil operating hours exceed 13.5 million. As a further development, a completely fluidic engine control system is proposed with an electronic computer controlled secondary control or trim system for efficiency optimization.

Optimizing the Architecture of Civil Turbofan Engines to Improve Life Cycle Costs/Value Added

2004 ◽  
Author(s):  
Marc Foerstemann ◽  
Stephan Staudacher
Keyword(s):  
Life Cycle ◽  
Value Chain ◽  
Life Cycle Cost ◽  
Product Life Cycle ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Value Added ◽  
Aircraft Engine ◽  
Life Cycle Costs ◽  
Turbofan Engines

In this paper, an economic calculation model is described, which evaluates the “life cycle value added” of an aircraft and an aircraft engine from the end customers’ — the airlines’ — perspective. The model uses both revenue as well as total costs over the entire product life cycle. It can be used to assess the economic benefit of a certain product (e.g., aircraft or engine), of a defined improvement measure or of different design options. Based on a complex set of parameters, the model can even be used in early design phases, where the potential impact on life cycle cost is the highest. The model is used to show that existing turbofan engines can be improved to deliver extra value for the end customer and as such for the entire value chain. Specific fuel consumption, manufacturing costs, maintenance costs, weight, drag and development costs are the most significant engine parameters for influencing the life cycle value added. An existing modern two-spool high bypass ratio engine was selected as the baseline configuration for applying the model. An analysis of the engine’s architecture identified the engine’s booster as a potential area of improvement. Upgrading the high-pressure compressor to the latest technology would enable the overall pressure ratio to be maintained while omitting the booster and improving engine performance. The results of the calculation show an improvement of life cycle value added, despite significant one-off development, testing and certification costs. The results support the hypothesis that today’s turbofan engines provide room for life cycle costs/value added improvement.

Impact of Air System Bleeding on Aircraft Engine Performance

2011 ◽  
Author(s):  
Bin Zhao ◽  
Shaobin Li ◽  
Qiushi Li ◽  
Sheng Zhou
Keyword(s):  
Fuel Consumption ◽  
Negative Impact ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Aircraft Engine ◽  
Engine Model ◽  
Engine Thrust ◽  
Air System ◽  
Compressor Performance

Air system bleeding is indispensable to aircraft engines despite its negative impact on the engine thrust and the fuel consumption. However, the compressor performance can be improved if the bleeding design is optimized. The model in this paper is a one-dimensional engine model based on air system bleeding. The relation between the compressor performance and the engine thermodynamic cycle caused by bleeding is analyzed to explore the potential of air system bleeding in improving compressor and engine performance. The results show that if bleeding brings an increase the pressure ratio of compressor, the negative impact on engine specific fuel consumption can be inhibited. If the efficiency of compressor is increased after bleeding, the negative impact on engine thrust can be alleviated. With proper bleeding flow rates, if both the pressure ratio and the efficiency increase at the same time, the negative impact on the engine performance can be eliminated.

scholarly journals Optimal Tuner Selection for Kalman Filter-Based Aircraft Engine Performance Estimation

2009 ◽  
Vol 132 (3) ◽  
Author(s):  
Donald L. Simon ◽  
Sanjay Garg
Keyword(s):  
Kalman Filter ◽  
Estimation Error ◽  
Engine Performance ◽  
Aircraft Engine ◽  
Operating Conditions ◽  
Performance Estimation ◽  
Experimental Simulation ◽  
Tuning Parameter ◽  
Estimation Accuracy ◽  
On Line

A linear point design methodology for minimizing the error in on-line Kalman filter-based aircraft engine performance estimation applications is presented. This technique specifically addresses the underdetermined estimation problem, where there are more unknown parameters than available sensor measurements. A systematic approach is applied to produce a model tuning parameter vector of appropriate dimension to enable estimation by a Kalman filter, while minimizing the estimation error in the parameters of interest. Tuning parameter selection is performed using a multivariable iterative search routine that seeks to minimize the theoretical mean-squared estimation error. This paper derives theoretical Kalman filter estimation error bias and variance values at steady-state operating conditions, and presents the tuner selection routine applied to minimize these values. Results from the application of the technique to an aircraft engine simulation are presented and compared with the conventional approach of tuner selection. Experimental simulation results are found to be in agreement with theoretical predictions. The new methodology is shown to yield a significant improvement in on-line engine performance estimation accuracy.

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