Landing

From the market requirements, Vapp = 120 knots = 120 x 1.68781 = 202.5 ft/s (61.72 m/s). Landing CLmax = 2.1 at a 40-deg flap setting (from testing and CFD analysis). For sizing purposes, the engine is set to the idle rating, producing zero thrust using Equation 11.22.

In the FPS system, W/Sw = 0.311 x 0.002378 x 2.1 x (202.5)2 = 63.8 lb/ft2. In the SI system, W/Sw = 0.311 x 1.225 x 2.1 x (61.72)2 = 3,052 N/m2. Because the thrust is zero (i. e., idle rating) at landing, the W/Sw remains constant.

Performance. Chapter 13 verifies whether the design meets the aircraft perfor­mance specifications.

11.3 Coursework Exercises: Military Aircraft Design (AJT)

This extended section of the book on coursework exercises – military aircraft design (AJT) is found on the Web at www. cambridge. org/Kundu and includes the following subsections.

11.5.1 Takeoff – Military Aircraft

Table 11.4. AJT takeoff sizing

11.5.2 Initial Climb – Military Aircraft

Table 11.5. AJT climb sizing

11.5.3 Cruise – Military Aircraft

Table 11.6. AJT cruise sizing

11.5.4 Landing – Military Aircraft

11.4 Sizing Analysis: Civil Aircraft (Bizjet)

The four sizing relationships (Sections 11.3.1 through 11.3.4) for wing-loading, W/Sw, and thrust-loading, TSlS_inStalled/W, meet (1) takeoff, (2) approach speed

Figure 11.3. Aircraft sizing: civil aircraft

for landing, (3) initial cruise speed, and (4) initial climb rate. These are plotted in Figure 11.3.

The circled point in Figure 11.3 is the most suitable for satisfying all four requirements simultaneously. To ensure performance, there is a tendency to use a slightly higher thrust-loading TSLS_INSTALLED/W; in this case, the choice becomes Tsls-installed/W = 0.32 at a wing-loading of W/SW = 63.75 lb/ft2 (2,885 N/m2).

Now is the time for the iterations for the preliminary configuration generated in Chapter 6 from statistics, in which only the fuselage was deterministic. At 20,720 lb (9,400 kg) MTOM, the wing planform area is 325 ft2, close to the original area of 323 ft2; hence, no iteration is required. Otherwise, it is necessary to revisit the empennage sizing and revise the weight estimates. The TSLSINSTALLED per engine then becomes 0.32 x 20,720/2 = 3,315 lbs. At a 7% installation loss at takeoff, this gives uninstalled TSLS = 3,315/0.93 = 3,560 lb/engine (TSLS/W = 3,560 x 2/20,720 = 0.344). This is very close to the TFE731-20 class of engine; therefore, the engine size and weight remain the same. For this reason, iteration is avoided; otherwise, it must be carried out to fine-tune the mass estimation.

The entire sizing exercise could have been conducted well in advance, even before a configuration was settled – if the chief designer’s past experience could “guesstimate” a close drag polar and mass. Statistical data of past designs are useful in guesstimating aircraft close to an existing design. Mass fractions as provided in Section 8.8 offer a rapid mass estimation method. Generating a drag polar requires some experience with extraction from statistical data.

In the industry, more considerations are addressed at this stage – for example, what type of variant design in the basic size can satisfy at least one larger and one smaller capacity (i. e., payload) size. Each design may have to be further varied for more refined variant designs.

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