Wittgenstein's combustion chamber

2009 ◽  
Vol 63 (1) ◽  
pp. 95-104 ◽  
Author(s):  
John Cater ◽  
Ian Lemco
Keyword(s):  
Combustion Chamber ◽  
Ludwig Wittgenstein ◽  
Western World ◽  
Propeller Blade ◽  
Engineering Research ◽  
Probable Function ◽  
Variable Volume ◽  
Aero Engine ◽  
Blade Tip ◽  
Aeronautical Engineering

Ludwig Wittgenstein, destined to be one of the most influential philosophers of the western world, entered Manchester University in 1908 as an aeronautical engineering research student. At Manchester he devised and patented a novel aero-engine that employed propeller-blade tip-jets. As a first practical step to the realization of this device, Wittgenstein constructed a variable-volume combustion chamber, but on departing for Cambridge he abandoned all further work on the project. The plans of this chamber survived and are presented in this paper. This article includes a detailed description of the drawings and an analysis of the probable function of the system.

Wittgenstein's aeronautical investigation

2006 ◽  
Vol 61 (1) ◽  
pp. 39-51 ◽  
Author(s):  
Ian Lemco
Keyword(s):  
Twentieth Century ◽  
Gas Turbine ◽  
Royal Society ◽  
Physical Sciences ◽  
Engineering Research ◽  
German Education ◽  
Aero Engine ◽  
Blade Tip ◽  
Aeronautical Engineering ◽  
Flow Compression

After a rigorous German education in the physical sciences, young Ludwig Wittgenstein entered Manchester University as an aeronautical engineering research student. There he devised and patented a novel aero-engine employing an airscrew propeller driven by blade tip-jets. Within the context of the growth of English aviation during the first half of the twentieth century (including the contributions of many Fellows of the Royal Society) and taking into account related aspects of his life, this paper examines an unfulfilled engineering aspiration. In enlarging upon what Wittgenstein might have accomplished during his stay at Manchester, it contrasts his invention with later comparable proven designs, albeit applied to hybrid rotorcraft. His engine employed centrifugal flow compression and arguably was a precursor of Sir Frank Whittle's gas turbine. In conclusion, reasons are given for Wittgenstein's departure from Manchester.

Numerical and Experimental Investigations on Liner Heat Transfer in an Aero Engine Combustion Chamber

2017 ◽  
Author(s):  
Korukonda Venkata Lakshmi Narayana Rao ◽  
B. V. S. S. S. Prasad ◽  
Ch. Kanna Babu ◽  
Girish K. Degaonkar
Keyword(s):  
Combustion Chamber ◽  
Structural Integrity ◽  
Operating Conditions ◽  
Experimental Investigations ◽  
Numerical Computations ◽  
Combustion Gases ◽  
Flow Characterization ◽  
Aero Engine ◽  
Straight Flow ◽  
Flow Variables

The Gas turbine combustion chamber is the highest thermally loaded component where the temperature of the combustion gases is higher than the melting point of the liner that confines the gases. Combustor liner temperatures have to be evaluated at all the operating conditions in the operating envelope to ensure a satisfactory liner life and structural integrity. On experimental side the combustion chamber rig testing involves a lot of time and is very expensive, while the numerical computations and simulations has to be validated with the experimental results. This paper is mainly based on the work carried out in validating the liner temperatures of a straight flow annular combustion chamber for an aero engine application. Limited experiments have been carried out by measuring the liner wall temperatures using k-type thermocouples along the liner axial length. The experiments on the combustion chamber testing are carried out at the engine level testing. The liner temperature which is numerically computed by CHT investigations using CFX code is verified with the experimental data. This helped in better understanding the flow characterization around and along the liner wall. The main flow variables used are the mass flow rate, temperature and the pressure at the combustor inlet. Initially, the fuel air ratio is used accordingly to maintain the same T4/T3 ratio. The effect of liner temperature with T3 is studied. Since T4 is constant, the liner temperature is only dependent on T3 and follows a specific temperature distribution for the given combustor geometry. Hence this approach will be very useful in estimating the liner temperatures at any given T3 for a given combustor geometry. Further the liner temperature is also estimated at other fuel air ratios (different T4/T3 ratios) by using the verified CHT numerical computations and found that TL/T3 remains almost constant for any air fuel ratio that is encountered in the operating envelope of the aero 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).

A Novel Generalized Predictive Control and its Application on Active Clearance Control of Aero-Engine

2012 ◽  
Vol 616-618 ◽  
pp. 1922-1925
Author(s):  
Kai Peng ◽  
Ding Fan ◽  
Lei Zhang ◽  
Qiu Xia Wang
Keyword(s):  
Predictive Control ◽  
Gas Turbines ◽  
Dynamic Performance ◽  
Tip Clearance ◽  
Generalized Predictive Control ◽  
Aero Engine ◽  
Blade Tip ◽  
Additional Service ◽  
And Control ◽  
Clearance Control

Turbine blade tip clearance continues to be a concern in the design and control of gas turbines. Ever increasing demands for improved efficiency and higher operating temperatures require more stringent tolerances on turbine tip clearance. An implicit active generalized predictive control with AR error modification and fuzzy adjustment on control horizon of aero-engine turbine tip clearance is presented and evaluated. The results show the resultant active tip clearance control system has good steady and dynamic performance and benefits of increased efficiency, reduced specific fuel consumption, and additional service life.

scholarly journals Flow field and species concentration measurements in the primary zone of an aero-engine combustion chamber

2018 ◽  
Vol 10 (1) ◽  
pp. 168781401774805
Author(s):  
Yinli Xiao ◽  
Zupeng Wang ◽  
Zhengxin Lai ◽  
Kefei Chen ◽  
Wenyan Song
Keyword(s):  
Flow Field ◽  
Combustion Chamber ◽  
Operating Conditions ◽  
Chamber Pressure ◽  
Species Concentration ◽  
Primary Zone ◽  
Major Species ◽  
Engine Combustion ◽  
Aero Engine ◽  
Concentration Measurements

The principal features of primary zone determine the performance parameters of the whole combustion chamber, such as the pollutant emissions and combustion efficiency. In this work, flow field and major species concentration measurements are conducted in the primary zone of an aero-engine combustion chamber. The operating conditions such as air inlet temperature, chamber pressure, and air-to-fuel ratio are chosen to replicate the realistic operating conditions. The velocity field and streamlines are obtained by particle imaging velocimetry technology. The concentrations of major species are acquired by a spontaneous Raman scattering system. This article validates the feasibility of two laser diagnostic measurement techniques and presents the initial results under realistic aero-engine conditions.

scholarly journals THE STUDY OF SPECIAL ASPECTS OF COMBUSTION IN A VARIABLE VOLUME COMBUSTION CHAMBER

2019 ◽  
pp. 64-71
Author(s):  
A. P. Shaikin ◽  
◽  
P. V. Ivashin ◽  
I. R. Galiev ◽  
I. N. Bobrovsky ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Variable Volume

Movement of Deposited Water on Turbomachinery Rotor Blade Surfaces

2006 ◽  
Author(s):  
John Williams ◽  
John B. Young
Keyword(s):  
Rotor Blade ◽  
Water Movement ◽  
Equations Of Motion ◽  
Surface Friction ◽  
Blade Surface ◽  
Inertia Effects ◽  
Aero Engine ◽  
Blade Tip ◽  
Fan Blade ◽  
Bypass Ratio

A theoretical approach for calculating the movement of liquid water following deposition onto a turbomachine rotor blade is described. Such a situation can occur during operation of an aero-engine in rain. The equation of motion of the deposited water is developed on an arbitrarily oriented plane triangular surface facet. By dividing the blade surface into a large number of facets and calculating the water trajectory over each one crossed in turn, the overall trajectory can be constructed. Apart from the centrifugal and Coriolis inertia effects, the forces acting on the water arise from the blade surface friction, and the aerodynamic shear and pressure gradient. Non-dimensionalisation of the equations of motion provides considerable insight and a detailed study of water flow on a flat rotating plate set at different stagger angles demonstrates the paramount importance of blade surface friction. The extreme cases of low and high blade friction are examined and it is concluded that the latter (which allows considerable mathematical generalisation) is the most likely in practice. It is also shown that the aerodynamic shear force, but not the pressure force, may influence the water motion. Calculations of water movement on a low-speed compressor blade and the fan blade of a high bypass ratio aero-engine suggest that in low rotational speed situations most of the deposited water is centrifuged rapidly to the blade tip region.

Combustion chamber design and reaction modeling for aero turbo-shaft engine

2018 ◽  
Vol 91 (1) ◽  
pp. 94-111
Author(s):  
Raja Marudhappan ◽  
Chandrasekhar Udayagiri ◽  
Koni Hemachandra Reddy
Keyword(s):  
Numerical Modeling ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Diffusion Flame ◽  
Internal Flow ◽  
Content Type ◽  
Combustion Modeling ◽  
Flame Combustion ◽  
Aero Engine ◽  
Exit Temperature

Purpose The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero turbo shaft engine. The purpose is extended to perform numerical combustion modeling by solving transient Favre Averaged Navier Stokes equations using realizable two equation k-e turbulence model and Discrete Ordinate radiation model. The presumed shape β-Probability Density Function (β-PDF) is used for turbulence chemistry interaction. The experiments are conducted on the real engine to validate the combustion chamber performance. Design/methodology/approach The combustor geometry is designed using the reference area method and semi-empirical correlations. The three dimensional combustor model is made using a commercial software. The numerical modeling of the combustion process is performed by following Eulerian approach. The functional testing of combustor was conducted to evaluate the performance. Findings The results obtained by the numerical modeling provide a detailed understanding of the combustor internal flow dynamics. The transient flame structures and streamline plots are presented. The velocity profiles obtained at different locations along the combustor by numerical modeling mostly go in-line with the previously published research works. The combustor exit temperature obtained by numerical modeling and experiment are found to be within the acceptable limit. These results form the basis of understanding the design procedure and opens-up avenues for further developments. Research limitations/implications Internal flow and combustion dynamics obtained from numerical simulation are not experimented owing to non-availability of adequate research facilities. Practical implications This study contributes toward the understanding of basic procedures and firsthand experience in the design aspects of combustors for aero-engine applications. This work also highlights one of the efficient, faster and economical aero gas turbine annular diffusion flame combustion chamber design and development. Originality/value The main novelty in this work is the incorporation of scoops in the dilution zone of the numerical model of combustion chamber to augment the effectiveness of cooling of combustion products to obtain the desired combustor exit temperature. The use of polyhedral cells for computational domain discretization in combustion modeling for aero engine application helps in achieving faster convergence and reliable predictions. The methodology and procedures presented in this work provide a basic understanding of the design aspects to the beginners working in the gas turbine combustors particularly meant for turbo shaft engines applications.

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