Gas Turbine Theory

Gas Turbine Theory
QUEEN MARY, UNIVERSITY OF LONDON DEPARTMENT OF ENGINEERING DEN-306 AIRCRAFT PROPULSION JET ENGINE PERFORMANCE EXERCISE Introduction This exercise is concerned with a simple jet engine and involves using the NASA EngineSimU program. The program accepts data about the engine such as compressor pressure ratio, turbine inlet temperature and the altitude and speed of the aircraft, in which the engine is installed. The fuel consumption and thrust from a jet engine depends upon engine design and the operating condition of the aircraft. The performance of a jet engine has to be estimated at an early stage for the flight conditions for which the aircraft will be used. Further background on this will be found in the textbook ‘Gas Turbine Theory’ by Cohen, Rogers and Saravanamuttoo. A typical output is shown in the form of a performance map on Fig.3.12. (In the fifth edition).

From: Cohen, Rogers and Saravanamuttoo ‘Gas Turbine Theory’

The calculations carried out by the CATU program: The CATU program accepts the altitude and use empirical equations to determine the corresponding ambient pressure and temperature. The aeroplane Mach number is used with the above to calculate the stagnation temperature and pressure. Next the stagnation pressure delivered by the intake is calculated using the efficiency of the intake. The pressure and temperature rise through the compressor is calculated and the combustion temperature is evaluated by means of a numerical scheme which is equivalent to using the Rogers & Mayhew ‘Combustion temperature rise – fuel air ratio chart’. Following the combustion chamber the turbine calculation is carried out, knowing that the power from the turbine has to be sufficient to drive the compressor and turbine mass flow is the compressor mass flow + the fuel flow. The exhaust products from the turbine are expanded through the propelling nozzle which is normally of the convergent type and the gross thrust from the nozzle is calculated. The above information is used, with the intake momentum drag, to calculate the specific fuel consumption, SFC or TSFC in the case of EngineSimU. This is the fuel flow rate divided by the net thrust.

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Using the NASA EnginSimU program To use this program go http://www.grc.nasa.gov/WWW/k-12/airplane/engsimu.html (make sure that Java is enabled on your web browser). Choose the turbojet option and set the program for metric units. (i) Record the variation of the specific fuel consumption (SFC) with the specific net thrust (SNT) for an altitude of 7000 meters, aircraft Mach number of 0.8, ideal intake and nozzle, efficiencies of 0.9 for the compressor and turbine, pressure loss of 0.03 in the combustion chamber and efficiency of one, while varying the compressor pressure ratio from 6 to 9 in steps of one and the combustion chamber maximum temperature from 1100K to 1400K in steps of 100K. The SNT is to be calculated by dividing the net thrust by the mass airflow rate. Plot SFC .vs. SNT for these conditions using Excel or similar, producing a graph similar to that on the first page. [30%] (ii) Carry out a hand calculation for one of the conditions of the above; assume the compressor and turbine efficiencies are isentropic. Compare your results of SFC and SNT to the results produced by the computer program. Explain the difference. [30%]

(iii) For three compressor pressure ratios and three combustion temperatures, vary the Mach number (at least three Mach numbers, one supersonic) and altitude (at least three altitudes, one above 11km), and plot the variation of SFC and SNT with the Mach number and altitude [20%] (iv) Discuss the effects of the variation of SFC and SNT with the Mach number and altitude. [20%] You are required to produce only a brief report that answers the questions set above. No need for introduction or table of contents. Include references in your report if needed.

EJ Avital January 2014

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