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Relight envelope of the combustor needs to be experimentally generated and established during the design and development of an aero gas turbine engine. Usually, during development stage of engine, compressor characteristics are not readily available at such low speeds and hence, it becomes difficult to specify the combustor inlet conditions such as pressure, temperature and Mach number during the engine light up studies. This paper compares the experimental test data generated on an annular combustor for windmill conditions during stand-alone mode and engine level tests under simulated flight conditions. The stand-alone combustor trials were conducted for the range of total pressure and temperature relevant to the flight altitude and Mach number range. During the engine level tests, combustor relight tests were conducted under simulated conditions (ISA+15) for altitudes ranging from 5.5 km to 10 km, flight Mach numbers in the range of 0.45 to 0.80. In this paper, effect of altitude and flight Mach number on the windmill spool speed, combustor pressure and temperature are studied.

The Hot End Technologies Directorate (HETD) of Gas Turbine Research Establishment (GTRE) has the mandate to design, development and delivery of airworthy combustor and afterburner modules for a military aero gas turbine engine. In order to meet the mandate, the directorate takes the overall responsibility of design to manufacture of the combustion systems. This paper addresses the challenges faced in the development of combustor module. A short annular combustor with air blast atomizer is incorporated in the engine and it is a very important equipment of a gas turbine engine, wherein the heat energy is added to get Turbine Inlet Temperature (TET). It comprises of a pre-diffuser, a dump diffuser, outer annulus, inner annulus and a flame tube. There has been a basic liner, which was used in earlier engines and there was a shortfall in terms of performance parameters — allowable profile and pattern factors. To improve the performance, in collaboration with the M/s Central Institute of Aviation Motors (CIAM), Moscow, Russia, the liner was redesigned [1]. The secondary holes were totally blocked, primary and dilution holes were altered and it was incorporated with a new dome with a modified curvature. A new air blast atomizer with a swirler having an outer and inner pintle was incorporated. The basic liner was incorporated with these modifications and making this dome out of the high temperature resistance nickel chromium alloy was challenging and it was realized. The liner assemblies incorporating all the welding details have been realized within the GTRE. The combustor system was tested for ground light up to 4.3 km. The light up time was of the order of 5 s. The pressure loss was of the order of 4.9% at a combustor inlet Mn of 0.30. The circumferential and radial pattern factor for the modified liner is of 0.36 and 0.14 respectively.