Influence of hot and cold zones in gas flow at the turbine disk inlet on a thermal loading of rotating blades

 

Lapotko V.M., Kukhtin J.P., Bljumin J.I. ZMKB "Progress", Zaporozhye, The Ukraine

 

Aerodynamic simulations of flow over turbine blades is necessary for their reliable design with allowance for thermal loading in real conditions of operation. In the present report, the developed method to simulate unsteady gas flows (in other words, the method of streamline observation) is employed for the analysis of influence of hot and cold zones in gas flow at the turbine inlet on the thermal loading of blades. In a layer of a variable thickness and at a variable radius, two-dimensional flow was considered, which is described by a system of integro-differential equations written in curvilinear coordinate system (j, m) bound with rotating surface of reference of the turbine. Turbulent flow is under investigation. It is simulated using one-parameter model of turbulence proposed by Kovazhny. Heat and mass transfer processes caused by viscous and turbulent transfer of a substance are simulated by sources and sinks of the mass, momentum and energy.

Parameter, which determines thermal loading of blades, is total temperature of gas in its relative motion at the inlet to the disk. This temperature can vary considerably in circumferential direction because of:

1) Flow deceleration (up to its bringing to rest) in boundary layers on vanes;

2) Injection of cooling air for vane cooling;

3) Effect of the engine's combustion chamber.
The experimental high-pressure turbine stage of D-27 engine, which was put under investigation, has 28 vanes and 64 blades, a rotational disk speed is 19630 1/min.

Obtained via calculations, which had been made prior to the investigation, were distributions of total temperature in relative motion at the inlet to the calculation area. The maximal surplus of temperature in the wake because its bringing to rest in a vane boundary layer reaches approximately to 10 % of the total temperature in the free-stream flow.

More essential circumferential non-uniformity of total temperature at the inlet to the turbine rotor causes injection of cooling air to cool the vanes. The existing system of holes was substituted in predictions by an equivalent system of slots. Due to the peculiarity of used method, each slot injecting the cooling air was reproduced with a separate stream of flow at a given temperature. Set for each slot were a width, a location at the profile and the direction of injection. Total pressure in slots was selected so that it can ensure acceptable quality of vane film cooling at the given total flow-rate of air equal to 7 % of the gas flow-rate in the main inter-blade pass.

However, the most significant non-uniformity in the total temperature profile is very often induced by the combustion chamber. To reproduce this effect, a flow over guide vanes with the specified inlet temperature non-uniformity was simulated. It was supposed that the hot gas zone is located in the middle between leading edges of two guide vanes at the center of each other stator channel and comprised about 50 % of a pitch of vane row. The total pressure was kept constant, while the temperature rise in the hot zone was reproduced by the reduction in the gas density. Investigated during simulations was the level of thermal non-uniformity with the ratio of maximum temperature to the free stream temperature equal to 1.5. Obtained in the research was also the total temperature distribution at the outlet from nozzle guide vanes at presence of above-mentioned temperature non-uniformity induced by the combustion chamber as well as air injection for guide vane cooling.

In order to carry out the analysis of the thermal state of blades, a relative flow field was calculated around the turbine disk sector consisting of 7 blades. This sector of blades was assumed to be stationary, whilst the wake non-uniformity at the left boundary of the calculated area became to be moving from the bottom in the upward direction with a circumferential velocity equal to the operational rotational speed.

Included in the set of parameters characterizing the thermal loading of blades were time-average (during passing through one period of wake) distributions of the total temperature in the circumferential direction at the cross-section drawn through leading edges of the investigated profiles as well as analogous time-average total temperature distributions over convex and concave surfaces of the blade.

According to obtained results, the circumferential non-uniformity of the temperature field in relative flow over turbine blades is preserved. The explanation of this phenomenon, which was named as segregation of the non-uniform gas flow over blade convex and concave surfaces, was given for the first time by American scientists. Thus an explanation has been given to the numerous experimental measurements made at the ZMKB "Progress" that showed higher gas temperatures on concave surfaces of blades in comparison with the gas temperatures on convex surfaces.

In calculations at turbine inlet temperature T = 1340°K, surplus gas temperature on the blade concave surface in relation to that on the blade convex surface at the middle part of a blade profile is equal to:
· approximately 10° because of the effect of the boundary layer on vanes;
· approximately 35° due to cooling air injection for vane cooling;
· approximately 120° due to the temperature non-uniformity induced by the combustion chamber.

The considered problem will be essentially complicated and require engaging additional computational tools, for example, in a case where temperature non-uniformity induced by the combustion chamber is non-periodic. Such a problem will require considering heat transfer and gas dynamics over the complete turbine stage. The resources of the used algorithms and modern computers allow fulfilling such analysis.