Maximum Speed Requirements (AJT)

An aircraft at HSC is at Mach 0.85 (845.5 ft/s) at a 30,000-ft altitude (p = 0.00088 slug/ft3). The fuel burned to climb is computed (but not shown) as 582 lb. The air­craft weight at the altitude is 10,000 lb.

At Mach 0.85, the aircraft lift coefficient CL = MTOM/gSV = 10,000/ (0.5 x 0.00088 x 845.52 x 183) = 10,000/57,561.4 = 0.174.

The clean aircraft drag coefficient (see Figure 9.16) at CL = 0.174 gives CDclean = 0.025 (high speed). The clean aircraft drag, D = 0.025 x (0.5 x 0.00088 x 858.52 x 183) = 0.025 x 5,7561.4 = 1,440 lb.

The available engine-installed thrust at the maximum cruise rating (i. e., 85% of the maximum rating) is from Figure 13.4 at Mach 0.85, and at a 30,000-ft altitude is T = 0.85 x 2,000 = 1,700 lb. (In the industry, the thrust is computed.)

Therefore, the AJT satisfies the customer requirement of Mach 0.85 at HSC.

13.6.2 Fuel Requirements (AJT)

Other than a ferry flight, military aircraft are not dictated only by the cruise sector, unlike in a civil aircraft mission. A short combat time at the maximum engine rating, mostly at low altitudes, is responsible for a suitable part of the fuel consumed. How­ever, the range to the target area dictates the fuel required. A long-distance ferry flight and combat arena require additional fuel to be carried by drop tanks. Imme­diately before combat, the drop tanks (they are empty) by then can be jettisoned to gain aircraft performance capability. The CAS variant has this type of mission profile.

A training mission has a varied engine demand and it returns to its own base covering no range, as shown in Figure 13.19. Mission fuel is computed sector by sec­tor of fuel burn, as shown as follows for the coursework example. To compute the fuel requirement, climb and specific-range graphs for the AJT at NTC are required (Figures 13.22 and 13.23). To compute the varied engine demand of a training – mission profile, Figure 13.4 is used to establish the fuel-flow rate for the throttle set­tings. The graph is valid for 75% rpm to 100% ratings. Typically, it has the approxi­mate following values:

• at idle (50% rpm) « 8 kg/min

• at 75% rpm « 11 kg/min

• at 95% rpm « 16.5 kg/min

Figure 13.22. AJT climb performance

Fuel and time consumed for the NTC of the AJT is shown in Table 13.20.

13.5 Summary

This chapter is the culmination of progress on the configuring, sizing, and substan­tiating of the coursework examples. It is time to review whether the Bizjet and the AJT designs need any revision. With commonality in design considerations, the tur­boprop aircraft is not addressed herein. The remaining chapters contain information on topics in which designers must be knowledgeable.

The sizing exercise (see Chapter 11) provides a simultaneous solution to satisfy airworthiness and market requirements. Wing-loading (W/SW) and thrust-loading (T/W) are the dictating parameters and they appear in the equations for takeoff, second-segment climb, enroute climb, and maximum speed capability; the first two are FAR requirements, the last two are customer requirements. Detailed informa­tion on engine performance is not required during the sizing exercise. Substantiation

Table 13.20. AJT mission fuel and time consumed

Fuel burned kg

Time

min

Engine rating = % rpm

Taxi and takeoff

60

6

60% (idle)

Takeoff and climb to 6-km altitude

125

5

Takeoff @ 100%, then @ 95%

Four turns

50

4

1 min @ 95% + 3 min @ 60%

Four stalls

60

5

1 min @ 95% + 4 min @ 60%

Climb from 5- to 8-km altitude

50

3

95%

Four turn spins

25

3

60%

Climb from 5- to 8-km altitude

50

3

95%

Four turn spins

25

3

60%

Climb from 5- to 6-km altitude

15

1

95%

Aerobatics practice

70

6

95%

Descent and practice force landing

95

8

2 min @ 95% + 6 min @ 60%

Three circuits for landing practice

110

10

Average 80%

Approach, land, return taxi

40

4

60%

Trainee pilot allowance

30

2

95%

Total mission fuel

805

59 (^60)

Diversion Residual fuel

Total onboard fuel

200

105

1,110 (conservative estimate)

(Internal fuel capacity =

1,400 kg)

of the payload-range estimation, as a customer requirement, is not possible dur­ing the sizing exercise beacuse it requires detailed engine performance data. Sub­sequently, with detailed engine performance data, relevant aircraft performance analyses are conducted more accurately to guarantee airworthiness and market requirements.

A more detailed aircraft performance is estimated during the Detailed Defini­tion Phase, which is beyond the scope of this book. The full aircraft performance does not affect aircraft configuration and mass unless the design review results in new demands for changes. These are management issues that are reviewed with potential customers to decide whether to give a go-ahead. Once a go-ahead is obtained, a full-blown detailed definition study ensues as Phase 2 activities, with significant financial commitments.

Figure 11.3 (Bizjet) and Figure 11.5 (AJT) show the lines of constraints for the various sizing requirements. The sizing point to satisfy all requirements shows a different level of margins for each capability. Typically, the initial enroute climb rate is the most critical to sizing. Therefore, the takeoff and maximum speed capabilities have a considerable margin, which is desirable because the aircraft can do better than what is required.

From statistics, experience shows that aircraft mass grows with time. This occurs primarily due to modifications resulting from mostly minor design changes and changing requirements – at times, even before the first delivery is made. If new requirements demand several changes, then a civil aircraft design may appear as a new variant. However, military aircraft design holds a little longer before a new variant emerges. It is therefore prudent for designers to maintain some margin, especially reserve thrust capability – that is, keep the thrust-loading (T/W) slightly higher. Re-engining with an updated version is costly.

It can be seen that field performance requires a larger wing planform area (SW) than at cruise. It is advisable to keep the wing area as small as possible (i. e., high wing-loading) by incorporating a superior high-lift capability, which is not only heavy but also expensive. Designers must seek a compromise to minimize operating costs (see Chapter 16). Iterations were not needed for the designs worked out in this book.

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