INSTALLATION LOSSES

The performance curves that have been presented for the JT4A-3 tur­bojet, the JT9D-7A turbofan, and the PT6A-27 turboprop are all optimistic, since they do not include installation losses. These losses result from:

• Total pressure loss in the inlet ducting.

• Total pressure loss in the exhaust nozzle.

• Bleed air requirements.

• Power extraction for accessories.

• Deicing requirements.

Methods for calculating these losses are not included here because of the extensive information that is required. In practice, an engine manufacturer supplies a computer deck to the airframe manufacturer in order to estimate corrections to the engine performance resulting from installation losses. Typically, these losses equal approximately 0.4% for inlet, 5% for antiicing, and 8 to 22 hp/engine for accessories.

TRENDS IN AIRCRAFT PROPULSION

The title of this section was borrowed directly from an interesting paper (Ref. 6.11) by Rosen. From the preceding material and examples of specific engines, it is hoped that you now have a pretty good feeling of engine performance capabilities in the 1978 time frame. Let us now consider what developments we might expect in the near future. Our considerations will be limited to subsonic airspeeds. To do otherwise is beyond the limitations of this text. It may also be beyond the price that society is willing to pay for speed with the emphasis on fuel economy and noise.

Regarding fuel consumption, Figure 6.47 presents the static TSFC as a function of net thrust for turbojets with afterburners, turbojets, and tur­bofans. The points represent engines that are currently operational. Generally,

there is a tendency, as with piston engines, for the TSFC to improve with size for any given engine type, particularly at the lowest thrust values. However, the important point is the obvious gain to be realized by going to higher bypass ratios. This point is emphasized by Rosen in a slightly different manner, as shown in Figure 6.48 (taken from Ref. 6.11).

* It is interesting that, in a sense, the application of gas turbines to commercial aircraft propulsion has nearly completed a cycle. The sudden transition to the turbojet for commercial transportation in the 1950s intro­duced the air traveler to above-the-weather flying at significantly higher speeds with a power plant that was almost vibrationless. In doing so, the bypass ratio went from a high value, where most of the air goes through the propulsor as compared to the air that goes through the power plant, to a value of zero, where all of the aif goes through the power plant. Over the years, the

BPR has gradually increased, but at no sacrifice in comfort or convenience to the passenger. Indeed, today’s high bypass ratio turbofan is quieter, consumes less fuel, and is relatively much lighter (Figure 6.49) than the turbojet.

As Rosen notes in Figures 6.48 and 6.49, for a given BPR, there is a gradual improvement in engine performance with time. This improvement is the result of better materials and cooling techniques, which allow operation at higher pressure ratios and turbine inlet temperatures. It would. therefore appear that improved propulsion efficiency in the future will depend on further increases in the pressure ratio, turbine inlet temperatures, and bypass ratios.

In 1971, Rosen was a fairly accurate soothsayer when he targeted turbine inlet temperatures of 2700 °F (1480 °С) and pressure ratios of 30:1 in the 1980s. We are not there yet, but the numbers are getting close. He also recommended a further increase in the bypass ratio by advocating a so-called prop-fan configuration. This configuration is a controllable pitch, ducted fan (or one might call it a propeller) with 8 to 12 blades, a 1.1 to 1.2 pressure ratio across the fan, and a tip speed of around 750 fps (230 m/s).

Figure 6.48 Specific fuel consumption trend. (George Rosen, “Trends in Aircraft Propulsion,” CASI Paper No. 72/10, 12th Anglo-American Aeronautical Con­ference, Canadian Aeronautics and Space Institute, 1971. Reprinted by per­mission of Canadian Aeronautics and Space Institute.)

Figure 6.49 Specific thrust trend. (George Rosen, “Trends in Aircraft Propul­sion,” CASI Paper No. 72/10, 12th Anglo-American Aeronautical Conference, Canadian Aeronautics and Space Institute, 1971. Reprinted by permission of Canadian Aeronautics and Space Institute.)

More recently, NASA has been taking another look at propellers for application to high subsonic speeds. Reference 6.12 discusses the design philosophy of these propellers and reports on some early test results. These propellers are multibladed and incorporate thin, transonic airfoil sections. One such propeller that has been tested by NASA’s Lewis Research Center is shown in Figure 6.50. Preliminary test data obtained in their supersonic wind tunnel and presented here as Figure 6.51 appear promising.

The design characteristics for this propeller are:

O. g cruise Mach number.

10.7 km (35,000 ft) cruise altitude.

8 blades.

203 activity factor per blade.

301 kW/m2 (37.5 hp/ft2) power loading. 243.8 m/s (800 fps) tip speed.

0.08 integrated design lift coefficient.

Mach number, M

Figure 6.51 Preliminary test data on advanced turboprop NASA Lewis Research Center 8- by 6-ft wind tunnel. A1.

If this propeller meets its design goal of 80% efficiency at 0.8 Mach

number, this would represent a reduction of approximately 30% in fuel

consumption compared to existing turbofan engines.