Page 157 - Fluid Mechanics and Thermodynamics of Turbomachinery

138 Fluid Mechanics, Thermodynamics of Turbomachinery
stall caused by a large adverse pressure gradient. So as to limit the total pressure
losses during flow diffusion it is necessary for the rate of deceleration (and turning)
in the blade passages to be severely restricted. (Details of these restrictions are
outlined in Chapter 3 in connection with the correlations of Lieblein and Howell.)
It is mainly because of these restrictions that axial compressors need to have many
stages for a given pressure ratio compared with an axial turbine which needs only a
few. Thus, the reversed turbine experiment tried by Parsons was doomed to a low
operating efficiency.
The performance of axial compressors depends upon their usage category.
Carchedi and Wood (1982) described the design and development of a single-shaft
15-stage axial-flow compressor which provided a 12 to 1 pressure ratio at a mass
flow of 27.3 kg/s for a 6 MW industrial gas turbine. The design was based on
subsonic flow and the compressor was fitted with variable stagger stator blades to
control the position of the low-speed surge line. In the field of aircraft gas turbines,
however, the engine designer is more concerned with maximising the work done
per stage while retaining an acceptable level of overall performance. Increased stage
loading almost inevitably leads to some aerodynamic constraint. This constraint will
be increased Mach number, possibly giving rise to shock-induced boundary layer
separation or increased losses arising from poor diffusion of the flow. Wennerstrom
(1990) has outlined the history of highly loaded axial-flow compressors with special
emphasis on the importance of reducing the number of stages and the ways that
improved performance can be achieved. Since about 1970 a significant and special
change occurred with respect to one design feature of the axial compressor and
that was the introduction of low aspect ratio blading. It was not at all obvious why
blading of large chord would produce any performance advantage, especially as the
trend was to try to make engines more compact and lighter by using high aspect
ratio blading. Wennerstrom (1989) has reviewed the increased usage of low aspect
ratio blading in aircraft axial-flow compressors and reported on the high loading
capability, high efficiency and good range obtained with this type of blading. One
early application was an axial-flow compressor that achieved a pressure ratio of
12.1 in only five stages, with an isentropic efficiency of 81.9% and an 11% stall
margin. The blade tip speed was 457 m/s and the flow rate per unit frontal area was
2
192.5 kg/s/m . It was reported that the mean aspect ratio ranged from a “high” of 1.2
in the first stage to less than 1.0 in the last three stages. A related later development
pursued by the US Air Force was an alternative inlet stage with a rotor mean aspect
ratio of 1.32 which produced, at design, a pressure ratio of 1.912 with an isentropic
efficiency of 85.4% and an 11% stall margin. A maximum efficiency of 90.9% was
obtained at a pressure ratio of 1.804 and lower rotational speed.
The flow within an axial-flow compressor is exceedingly complex which is one
reason why research and development on compressors has proliferated over the
years. In the following pages a very simplified and basic study is made of this
compressor so that the student can grasp the essentials.

Two-dimensional analysis of the compressor stage

The analysis in this chapter is simplified (as it was for axial turbines) by assuming
the flow is two-dimensional. This approach can be justified if the blade height is