Basic Research of the Intercooled Turbofan Aero-Engine

2012 ◽  
Vol 516-517 ◽  
pp. 544-547
Author(s):  
Jin Chuan Zhang ◽  
Yun Wang ◽  
Can Zhang
Keyword(s):  
Pressure Ratio ◽  
Basic Research ◽  
Inlet Temperature ◽  
Outlet Temperature ◽  
Gas Temperature ◽  
High Pressure Compressor ◽  
Aero Engine ◽  
Aero Engines ◽  
Cooling Air ◽  
Aircraft Application

In conventional turbofan aero-engine designs, the effective way of improvement of engine efficiency is through the increasing of overall pressure ratio and improving of combustor inlet gas temperature, but the further incresement of compressor overall pressor ratio is constricted by high pressure compressor outlet allowed temperature. The improvement of combustor outlet temperature is limited by turbine allowed inlet temperature during take-off and climbing. An intercooled core can be designed with a significantly higher overall pressure ratio also with reduced cooling air requirements, providing a higher thermal efficiency compared with a conventional core. Through the basic analysis of performance of intercooler aeroengines. It indicated that the intercooled aero-engines can decrese the feul consume clearly and have a further potential in future civil aircraft application.

scholarly journals Performance analysis of an aero engine with inter-stage turbine burner

2017 ◽  
Vol 121 (1245) ◽  
pp. 1605-1626 ◽  
Author(s):  
F. Yin ◽  
A. Gangoli Rao
Keyword(s):  
Fuel Consumption ◽  
Thermal Efficiency ◽  
Pressure Ratio ◽  
Future Generation ◽  
Nox Emission ◽  
Inlet Temperature ◽  
Performance And Emission ◽  
Aero Engine ◽  
Aero Engines ◽  
Bypass Ratio

ABSTRACTThe historical trends of reduction in fuel consumption and emissions from aero engines have been mainly due to the improvement in the thermal efficiency, propulsive efficiency and combustion technology. The engine Overall Pressure Ratio (OPR) and Turbine Inlet Temperature (TIT) are being increased in the pursuit of increasing the engine thermal efficiency. However, this has an adverse effect on engine NOx emission. The current paper investigates a possible solution to overcome this problem for future generation Very High Bypass Ratio (VHBR)/Ultra High Bypass Ratio (UHBR) aero-engines in the form of an Inter-stage Turbine Burner (ITB). The ITB concept is investigated on a next generation baseline VHBR aero engine to evaluate its effect on the engine performance and emission characteristics for different ITB energy fractions. It is found that the ITB can reduce the bleed air required for cooling the HPT substantially (around 80%) and also reduce the NOx emission significantly (>30%) without penalising the engine specific fuel consumption.

scholarly journals Development of a Flexible Turbine Cooling Prediction Tool for Preliminary Design of Gas Turbines

2018 ◽  
Vol 140 (9) ◽  
Author(s):  
Feijia Yin ◽  
Floris S. Tiemstra ◽  
Arvind G. Rao
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Specific Power ◽  
Inlet Temperature ◽  
Turbine Cooling ◽  
Engine Efficiency ◽  
Semi Empirical ◽  
Cooling Air

As the overall pressure ratio (OPR) and turbine inlet temperature (TIT) of modern gas turbines are constantly being increased in the pursuit of increasing efficiency and specific power, the effect of bleed cooling air on the engine performance is increasingly becoming important. During the thermodynamic cycle analysis and optimization phase, the cooling bleed air requirement is either neglected or is modeled by simplified correlations, which can lead to erroneous results. In this current research, a physics-based turbine cooling prediction model, based on semi-empirical correlations for heat transfer and pressure drop, is developed and verified with turbine cooling data available in the open literature. Based on the validated model, a parametric analysis is performed to understand the variation of turbine cooling requirement with variation in TIT and OPR of future advanced engine cycles. It is found that the existing method of calculating turbine cooling air mass flow with simplified correlation underpredicts the amount of turbine cooling air for higher OPR and TIT, thus overpredicting the estimated engine efficiency.

Research and Development of 300kW Class Ceramic Gas Turbine (Development of the Static Ceramic Components for CGT303)

1995 ◽  
Author(s):  
Mitsuharu Murota ◽  
Issei Ohhashi ◽  
Yoshiyuki Ito ◽  
Sadao Arakawa
Keyword(s):  
Research And Development ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Important Task ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Turbine Inlet Temperature ◽  
Engine Operation ◽  
Heat Cycle ◽  
Ceramic Components

As the result of setting the low pressure ratio at 4.5, sizes of the static ceramic components forming the gas passage in CGT303 have been increased, and establishing reliability of these components was thought to be the most important task. So, the heat-cycle tests were conducted, in advance of the engine operation, and improvements have been made on their material and constructions. After conducting 600 times of the heat-cycle tests, so far, up to the gas temperature of 1200°C, we have succeeded in the engine operation at the turbine inlet temperature of 1200°C Examples of the problems encountered in the test and of the solutions therefore are introduced in this paper.

Disk Heat Transfer Analysis in a Heated Rotating Cavity With an Axial Throughflow

2012 ◽  
Author(s):  
Shuqing Tian ◽  
Yatao Zhu
Keyword(s):  
Heat Transfer ◽  
Boundary Condition ◽  
Rotating Disk ◽  
Outer Wall ◽  
Heat Transfer Analysis ◽  
Outlet Temperature ◽  
Rotating Cavity ◽  
Isothermal Boundary ◽  
Aero Engines ◽  
Cooling Air

In the rotating disk cavities of aero-engine compressors, buoyancy-induced flow and heat transfer can occur due to thermal gradients between cooling air and hot surfaces. The simplified rotating cavity with two plane discs, a shaft and a cylindrical rim has been investigated numerically and compared with the available measurements. Two models have been solved using a commercial CFD code, Fluent, with the RNG k-ε turbulence model. The first one is the conventional model with only fluid region solved, a temperature profile with the linear radial gradient imposed at the disk walls, and an isothermal boundary condition imposed at the shroud wall. The second one is the model with thick-walled disks and shroud, an adiabatic boundary condition imposed at the outer walls of the disks, and an isothermal boundary condition imposed at the outer wall of the shroud. The fluid and solid are coupled solved simultaneously. The disk temperatures are computed. In the present work, the numerical results are in reasonable agreement with the measurements. The computed disk temperatures in the second model have approximately linear radial gradients over the first three-quarters of the disks, and in the last quarter of the disks the temperature radial gradients are obviously non-linear. The different disk temperature profiles in these two models do not lead to obviously different disk heat transfers. The heat transfer in the rotating cavity leads to a considerable temperature increase of the cavity core fluid, therefore a corresponding increase of the outlet temperature. These two temperature increases are critical for the cooling design in aero-engines.

scholarly journals Development of the AGT101 Regenerator Seals

1987 ◽  
Author(s):  
C. A. Fucinari ◽  
J. K. Vallance ◽  
C. J. Rahnke
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Dynamic Test ◽  
Development Program ◽  
Analytical Techniques ◽  
Gas Turbine Engine ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Turbine Engine ◽  
Seal Design

The design and development of the regenerator seals used in the AGT101 gas turbine engine are described in this paper. The all ceramic AGT101 gas turbine engine was designed for 100 hp at 5:1 pressure ratio with 2500F (1371C) turbine inlet temperature. Six distinct phases of seal design were investigated experimentally and analytically to develop the final design. Static and dynamic test rig results obtained during the seal development program are presented. In addition, analytical techniques are described. The program objectives of reduced seal leakage, without additional diaphragm cooling, to 3.6% of total engine airflow and higher seal operating temperature resulting from the 2000F (1093C) inlet exhaust gas temperature were met.

Cross-Correlations and Structures of Aero-Engine Gas Path System Based on DCCA Coefficient and Rooted Tree

2015 ◽  
Vol 14 (02) ◽  
pp. 1550014 ◽  
Author(s):  
Keqiang Dong ◽  
Jie Fan ◽  
You Gao
Keyword(s):  
Complex System ◽  
Cross Correlation ◽  
Pressure Ratio ◽  
Rooted Tree ◽  
Coefficient Matrix ◽  
Rotor Speed ◽  
Gas Temperature ◽  
Cross Correlation Analysis ◽  
Aero Engine ◽  
Path System

Identifying the mutual interaction is a crucial problem that facilitates the understanding of emerging structures in complex system. We here focus on aero-engine dynamic as an example of complex system. By applying the detrended cross-correlation analysis (DCCA) coefficient method to aero-engine gas path system, we find that the low-spool rotor speed (N1) and high-spool rotor speed (N2) fluctuation series exhibit cross-correlation characteristic. Further, we employ detrended cross-correlation coefficient matrix and rooted tree to investigate the mutual interactions of other gas path variables. The results can infer that the exhaust gas temperature (EGT), N1, N2, fuel flow (WF) and engine pressure ratio (EPR) are main gas path parameters.

A Methodology for Measuring Turbocharger Adiabatic Maps in a Gas-Stand and its Usage for Calibrating Control Oriented and 1D Models at Early ICE Design Stages

2019 ◽  
Author(s):  
José Ramón Serrano ◽  
Francisco José Arnau ◽  
Luis Miguel García-Cuevas ◽  
Alejandro Gómez-Vilanova ◽  
Stephane Guilain ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Internal Combustion Engines ◽  
Gas Flow ◽  
System Analysis ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Boost Pressure ◽  
Gas Temperature ◽  
Hot Gas ◽  
The Impact

Abstract Turbocharged engines are the standard architecture for designing efficient spark ignition and compression ignition reciprocating internal combustion engines (ICE). Turbochargers characterization and modeling are basic tasks for the analysis and prediction of the whole engine system performance and this information is needed in quite early stages of the engine design. Turbocharger characteristics (efficiency, pressure ratio, mass flow rates...) traditionally rely in maps of pseudo non-dimensional variables called reduced variables. These maps must be used by reciprocating ICE designer and modeler not only for benchmarking of the turbocharger, but for a multiplicity of purposes, i.e: assessing engine back-pressure, boost pressure, load transient response, after-treatment inlet temperature, intercooler inlet temperature, low pressure EGR temperature, ... Maps of reduced variables are measured in gas-stands with steady flow but non-standardized fluids conditioning; neither temperatures nor flows. In concrete: turbine inlet gas temperature; lubrication-oil flow and temperature; water-cooling flow and turbo-machinery external heat transfer are non-standardized variables which have a big impact in assessing said multiplicity of purposes. Moreover, adiabatic efficiency, heat losses and friction losses are important data, hidden in the maps of reduced variables, which depend on the testing conditions as much as on the auxiliary fluids temperature and flow rate. In this work it is proposed a methodology to standardize turbochargers testing based in measuring the maps twice: in close to adiabatic and in diathermal conditions. Along the paper it is discussed with special detail the impact of the procedure followed to achieve said quasi-adiabatic conditions in both the energy balance of the turbocharger and the testing complexity. As a conclusion, the paper proposes a methodology which combines quasi-adiabatic tests (cold and hot gas flow) with diathermal tests (hot gas flow) in order to extract from a turbocharger gas-stand all information needed by engine designers interested in controlling or 1D-modelling the ICE. The methodology is completed with a guide for calibrating said control-oriented turbocharger models in order to separate aerodynamic efficiency (adiabatic) from heat transfer losses and from friction losses in the analysis of the turbocharger performance. The outsourced calibration of the turbocharger model allows avoiding uncertainties in the global ICE model calibration, what is very interesting for turbochargers benchmarking at early ICE-turbo matching stages or for global system analysis at early control design stages.

scholarly journals New Diagram and Equation for Thermodynamic Calculations for Gas Turbines

1984 ◽  
Author(s):  
B. Herrmann
Keyword(s):  
Gas Turbines ◽  
Liquid Fuel ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Iso Standard ◽  
Compressor Pressure Ratio ◽  
German Standard ◽  
Exhaust Gas Temperature ◽  
Acceptance Tests

On basis of ISO-Standard 2314, the German Standard Organisation (DIN) has prepared the German Standard DIN 4341, which deals with acceptance tests for gas turbines. Sample calculations have been included. In connection with the development of the sample calculations a new diagram for thermodynamic properties of air and products of combustion was developed on basis of -humid air as per ISO standard 2314 -standard gaseous fuel -standard liquid fuel This diagram allows exact calculation of performance data. Further, a simplified but relatively acurate formula is presented for calculating the turbine inlet temperature on basis of -compressor pressure ratio -exhaust gas temperature -thermal efficiency Development and limitation of this formula is presented.

Potential of reducing the environmental impact of aviation by using hydrogen Part II: Aero gas turbine design

2006 ◽  
Vol 110 (1110) ◽  
pp. 541-552
Keyword(s):  
Environmental Impact ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Point Of View ◽  
Outlet Temperature ◽  
Hydrogen Fuel ◽  
Technical Point ◽  
Engine Design ◽  
Aero Engine ◽  
Turbine Design

Abstract The main objective of the paper is to evaluate the potential of reducing the environmental impact of civil subsonic aviation by using hydrogen fuel. The paper is divided into three parts of which this is Part II. In Part I the background, prospects and challenges of introducing an alternative fuel in aviation were outlined. In this paper, Part II, the aero engine design when using hydrogen is covered. The subjects of optimum cruising altitude and airport implications of introducing liquid hydrogen-fuelled aircraft are raised in Part III. The study shows that burning hydrogen in an aero gas turbine seems to be feasible from a technical point of view. If the priority is to lower the mission fuel consumption, the results indicate that an engine employing increased combustor outlet temperature, overall pressure ratio and by-pass ratio, seems to be the most attractive choice. The mission NOx emissions, on the other hand, seem to be reduced by using engines with a weak core and lowered by-pass ratio.

Single vs. Two Stage High Pressure Turbine Design of Modern Aero Engines

1999 ◽  
Author(s):  
Fathi Ahmad ◽  
Alexander V. Mirzamoghadam
Keyword(s):  
High Reliability ◽  
Pressure Ratio ◽  
Weight Ratio ◽  
Single Stage ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Two Stage ◽  
Cost Weight ◽  
Advantages And Disadvantages ◽  
Secondary Air

In this paper, the two-stage shrouded HPT engine configuration rated at 22000 lbs thrust is used as the baseline from which a single stage HPT unshrouded design is systematically derived to evaluate the potential weight and cost advantage. The baseline thermodynamic cycle at the rated thrust level was modified in order to optimize the turbine inlet temperature, overall pressure ratio, and core flow with a single stage HPT and deliver competitive performance. The comparative study, although preliminary in depth, has led to the advantages and disadvantages associated with an unshrouded single versus a two-stage shrouded HPT design. The results compare design configuration, secondary air system, weight, safety, life, specific fuel consumption (SFC), and future thrust growth capability. The main advantages of the single stage application are reductions in cost and complexity of design, lower turbine gas temperature, and ease of maintenance. The main disadvantages are in reduced turbine polytropic/isentropic efficiency for HPC pressure ratio greater than 9, increased SFC, higher rim speed, higher HPT exit Mach number, higher bypass ratio to achieve the desired thrust level, and possibly higher weight. A quantitative statement on the reduction of engine cost/weight is premature until a detailed design and the associated cost-benefit is performed. The paper concludes by recommending that the design philosophy of the modern unmixed turbofan engine (single or two-stage HPT) leads to a balance between the selected turbine gas temperature versus the by-pass ratio in order to minimize cost and maximize the thrust-to-weight ratio and the cycle efficiency. In either ease, the expected high reliability and reduced engine cost/weight in the context of future thrust-growth capability need to be demonstrated by proven technology which seem to favor the two-stage HPT configuration.

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