Cemento

Multi-spool, Variable Geometry and Bleed For bett.

performance, the rotational speeds of the fan and com-pressor usually need to be different. In general the h;

pressure compressor (HPC) runs about twice as fast.

the fan. This is accomplished by attaching the compre sor and fan to different spools or shafts which run c(j centric to each other. Thus, in the dual spc configuration shown in Figure 3.28, the HPC is coi nected to the high pressure turbine by an outer spool ar.

the fan/booster are connected to the low pressure turbi by an inner spool. In some configurations, three concei trie spools are used.

Variable stators and inlet guide vanes are construct so that the stagger angle of the vane can be varied in

Counter-Rotating UDF Blades

Exhaust Nozzle

_J11 Ms'-I,

w

Stationary Support Structure

Figure 3.27 UDF™ Counter Rotation Schematic

3-30 FANS AND COMPRESSOl

Two Spool Compression

viov/vsv

»t*rtf«

• i M d l H W ) / * HPC

HPC - High Pressure Compressor IPC - Intermediate Pressure Compressor

Figure 3.28 Multi-Spool Configurations

controlled fashion. This allows the vane to be re-aligned to the changing air angles that occur as the operating condition of the compressor or fan changes. An ex-panded view of a variable stator system is shown in Fig-ure 3.29. The variable vane is rotated by means of the lever arm. These lever arms are attached to a ring on the outside of the compressor case Figure 3.30.

An example of a vector diagram showing the change in relative air angle and relative Mach number into the first stage rotor for two levels of preswiri is also presented in Figure 3.29. A variation in swirl from 0 to aps produces thechiajrigein W and.i(3.showR,

Bleed ports are shown in Figures 3.28 and 3.33. Bleed air is used for customer purposes, for turbine cooling, for active clearance control and in engine starting.

The need for variable geometry, multi-spools, bleed, or some combination of them can be seen by considering the engine-start situation. At low rotational speeds the compressor cannot put as much work into the air as it can at high speed, and consequently cannot compress the air as much. The fixed area of die rear stages then limits the amount of air that can be pumped. The front stages try to pump the flow and in the process are

back-Lever Arm Bushing

Shim

Case

•inr

C S

¥

-Lock Nut -Washer

"^f" ^^- Bushing Actuation Ring

- Inner Washer

•*. Vane

Shroud

D

iimiiiiniiiiii

Bushing

Seal

Effect of Variable IGV on Rotor Inlet W&£

No Preswiri

Preswiri

Teeth

Figure 3.29 Variable Stators and IGV's

FANS AND COMPRESSORS 3-31

Figure 3.30 Compressor Rotor and Case

pressured trying to force the less-dense air through the rear stages. Mid-stage bleed allows this excess mass flow to be removed and the front stages to pump. As speed increases, work input increases, the air is com-pressed and goes through the area of the rear stages. At this point, bleed is no longer required. Variable geome-try can also be used to control pumping in the front block. The higher the pressure ratio across a given block of stages on the same spool, the more severe the prob-lem.

Leakage in Compressors Aircraft gas turbine engines have a large number of locations where air leakage can occur and the cumulative effect of leakage on engine power, thrust and fuel usage can be significant. Typical leakage paths are shown in Figure 3.34. In order to pre-serve efficiency in modern engines, which can have very high overall pressure ratios in the compression system (e.g. 38:1), close attention must be paid to leakage, seals and clearance. To compound the problem, evidence from airline reports notes that a typical high bypass ratio engine has an SFC increase of about 1 to 1 1/2% per year. Periodic overhauls do not fully recover this effi-ciency loss. The final result can be an engine with an

SFC 3 to 10% higher than that of a new engine. Sealing clearances also have a significant effect on compressor stall margin and are directly responsible for thrust droop.

Primary gas path seals perform two functions. They minimize gas recirculation between the blade tips and the wail for unshrouded airfoils, between shroud seals and labyrinth teeth, and across blade dovetail rotors and platforms. Primary gas path seals also minimize gas leakage out of the primary gas path (across flanges, vari-able vane pivots, and compressor end seals).

The leakage of gas across the airfoil tip from the pres-sure surface to the suction surface and the subsequent in-teraction with the endwall boundary layer of the primary gas stream and the production of secondary flows can produce substantial loss in efficiency and stall margin. In one research compressor data were obtained at two lev-els of tip-clearance-to-blade-height ratio. 1.38% and 2.8%. This increase in tip clearance costs 1.5 points in peak efficiency, 11.0% in stalling flow coefficient (flow range) and 9.7% in peak pressure rise relative to the nominal clearance.

3-32 FANS AND COMPRESSORS

Consequently, designers go to great length to minimize leakage effects. An advanced compressor shown in ures 3.30 through 3.35 illustrates these measures. Fig-ure 3.30 shows a compressor with the upper half of the split casing removed. The rotor blading and the laby-rinth seal teeth are visible. The lower half of the casing shows the linkage mechanism for the variable stators and several bleed ports. Figure 3.31 shows the upper half of the casing. The stator vanes, the stator shrouds, and the shroud seal rub strips are visible. Figure 3.32 shows the IGVs, rotors, stators, stator shrouds, seal, and seal teeth.

Clearance Control The need for tight clearance from an aerodynamic perspective has already been established.

One way to achieve this is through active clearance con-trol. The variation in clearance between the casing and the rotor during engine operation is shown in Figure 3.36. The transient thermal response of the casing and the rotor and the centrifugal loading of the rotor are two of the most important factors in setting the final cruise clearance.

In general, the thermal response of the casing and the ro-tor are not the same because of differences in mass, cooling-air circulation, heat transfer, and material. The case tends to have a much faster thermal response to the gas stream temperature than the rotor. The rotor growth is initially due to centrifugal force during acceleration. If assembly clearances are too small, a rub will occur in the initial part of the acceleration due to the centrifugal force effect (Figure 3.36). On deceleration, the relatively fast thermal response of the casing will cause rubs if full power is demanded after a period of low power, such as an aborted landing. This happens because the relatively fast case response has reduced clearance to a magnitude less than the rotor displacement due to centrifugal force.

If very close cruise clearances are going to be main-tained, then some type of clearance control is needed to eliminate the rub potential. This can be accomplished by using bleed air to control casing and rotor temperatures, and thereby clearances. The compressor shown in Fig-ure 3.32 uses bleed air to control casing temperatFig-ures (diameter) and rotor diameter (bore cooling).

Summary of Aerodynamic Design Considerations Advances in technology over the past 20 years have led to substantial reductions in size, compactness, and weight in the modern compression system. In addition, significant improvements in efficiency have been ob-tained. These improvements can be seen (Figure 3.37) in the comparison of the engine cross-section for the E3

engine and the CF6-50 scaled to the same thrust.

AIRFOIL PHYSICAL AND FUNCTIONAL