Aircraft design projects - part 7 ppsx

“chap08” — 2003/3/10 — page 208 — #7 208 Aircraft Design Projects demanded of a strike aircraft. Good pilot visibility is also an advantage for the landing. Systems, including artificial vision and computer controlled imagery, will offer scope for innovation to overcome this problem in an aircraft designed for 2020. This aspect of layout and systems integration will require careful consideration. 8.3 Problem definition The project description specifies a two-place advanced deep interdictor aircraft. The entire long-range mission will beflown at supersonic speed. The exact missiondefinition is shown in Figure 8.1. The long-duration, high-intensity flight conditions, much of which is over enemy territory, demands the security of twin-pilot operation. The long work periods and high manoeuvre load environment imposed on the pilots requires careful design of the cockpit. The workload related to flight safety and weapon delivery must be reduced by system design. Such systems must be made reliable and safe. 1 2 3 4 5 6 7 8 9 10 11 12 Take-off Climb Climb Dash (in) Dash (return) Descend Descend Land (with reserves) Manoeuvre turn Supercruise (out) Supercruise (return) 1–2 Warm-up, taxi and take-off Sea level NATO 8000 ft, icy 2–3 Climb to best supercruise alt. 3–4 Supercruise to conflict area Opt. alt. M1.6 1000 nm 4–5 Climb to 50 000ft 5–6 Dash to target 50 000ft M1.6 750 nm 6–7 Turn and weapon release 50 000 ft 180° 7–8 Dash out 50 000ft M1.6 750 nm 8–9 Descend to supercruise alt. 9–10 Supercruise return 50 000ft M1.6 1000 nm 10–11 Descend to base 11–12 Land (with reserve fuel*) NATO 8000 ft, icy Segment Description Height Speed Distance/duration *Diversion and hold at sea level with 30 min fuel at economical fli g ht conditions. Fig. 8.1 Mission profile “chap08” — 2003/3/10 — page 209 — #8 Project study: advanced deep interdiction aircraft 209 The aircraft must be capable of ‘all-weather’ operation from advanced NATO and other bases. Aircraft shelter dimensions may impose configurational constraints on the aircraft. Aircraft servicing and maintenance at austere operational bases demand minimum support equipment and skill. Easy access to primary system components must be provided. Closed-loop, static and dynamic stability and handling flight characteristics must meet established military requirements. A digital flight control system will be necessary for a longitudinal unstable aircraft configuration. All systems must be protected against hostile damage and inherent unreliability. In addition to strict stealth criteria, the AIAA problem description sets out several required design capabilities and characteristics. These include: • The aircraft must accommodate two pilots but should be capable of single pilot operation. For such a long-range mission, pilot workload must be reduced by suitable design and specification of flight control and weapon delivery systems. Crew safety systems must be effective in all flight modes. • The design layout should allow for easy maintenance. Minimum reliance on support equipment is essential for off-base operations. • Structural design limit load factors of +7to−3g (aircraft clean and with 50 per cent internal fuel) are required. An ultimate design factor of 1.5 is to be applied. The structure must be capable of withstanding a dynamic pressure (q ) of 2133 lb/sq. ft (i.e. equivalent to (q) at 800 kt) and be durable and damage tolerant. • All fuel tanks must be self-sealing. Aviation fuel to JP8 specification (6.8 lb/US gal) is to be assumed. • Stability and handling characteristics to meet MIL-F-8785B subsonic longitudinal static margins to be no greater than +10 per cent and no less than −30 per cent. • The aircraft must be ‘all-weather’ capable. This includes operation from and on to icy 8000 ft runways. • The aircraft must operate from austere bases with minimum support facilities. On these bases the aircraft will be required to fit into standard NATO shelters. • The flyaway cost for 200 aircraft purchase must not exceed $150 M (year 2000 dollars). In addition to the high-altitude, supercruising mission shown in Figure 8.1 and described in section 8.2 above, the design specification sets the following manoeuvring targets (specific excess power, SEP, is defined as P S in Chapter 2 (section 2.7.1)): • SEP (1g) military thrust (dry), 1.6 M at 50 000 ft = 0 ft/s. • SEP (1g) maximum thrust (wet), 1.6 M at 50 000 ft = 200 ft/s. • SEP (2g) maximum thrust (wet), 1.6 M at 50 000 ft = 0 ft/s. • Maximum instantaneous turn rate, 0.9 M at 15 000 ft = 8.0 ◦ /s. (all the above performance criteria are specified at aircraft manoeuvre weight (defined as 50 per cent internal fuel with two AIM-120 and four 2000 lb JDAM)). The design specification calls for five separate weapon capabilities: • Four Mk-84 LDGP + two AIM-120. • Four GBU-27 + two AIM-120. • Four 2000 lb JDAM + two AIM-120. • Four AGM-154 JSOW + two AIM-120. • Sixteen 250 lb small smart bombs. “chap08” — 2003/3/10 — page 210 — #9 210 Aircraft Design Projects (the AIAA specification gave details of the size, weight and cost of all government furnished equipment. This data is used in the layout, mass and cost estimations). When details like those shown above are not provided with the initial specification, it is always necessary to spend time gathering the data before moving on to the next stage. In this case, we are now ready to consider initial aircraft design concepts. The details below suggest several potential design requirements: • The field take-off requirement, particularly with regard to the icy runway conditions will require a high thrust/weight ratio. • Initial climb performance will require good specific excess power to reach the supercruise altitude and speed in reasonable time. • Supercruise will require low overall drag to give a good lift/drag ratio and thereby a lower fuel requirement. • The rear movement of the centre of lift in supersonic flight may require fuel transfer to balance the aircraft and reduce trim drag. • The climb from supercruise altitude to 50 000 ft for the dash phase may require a burst of afterburning to offset the low SEP at high/fast operation. Stealth may be compromised by either the use of afterburning or from the long-duration climb from supercruise altitude to dash without the extra thrust. • The aircraft must be able to drop the weapons without significant trim changes. • The SEP requirements and the turn performance may require the use of manoeuvring flaps although this may compromise stealth. • Landing will require low wing loading to avoid high approach speed and to reduce aircraft energy on the ground. • Icy conditions may demand aerodynamic braking assistance (parachutes and lift dumping). • Compatibility with NATO shelter size will limit the aircraft to a span of less than 20 m (65 ft) and length to less than 30 m (98 ft). 8.4 Design concepts and selection Although initially many design layouts were envisaged, the three design concepts described below were selected for investigation. • Conventional, straight wing • Pure delta/diamond • Blended delta The conventional, tapered-wing layout (Figure 8.2) was selected as this offers less technical risk to the project. The design processes for this layout are well understood and the configuration can be easily developed for alternative roles. The pure arrow-wing layout (Figure 8.3) results from considerations of stealth and aerodynamic efficiency. The main drawbacks of the diamond planfor m centre on the unorthodox control arrangement and the difficulty of developing the layout to accommodate alternative roles. The blended arrow-wing configuration (Figure 8.4) can be regarded as either offering the best of the other options, or the worst of both types! The blended body can be configured to give lower wave drag than the straight wing and could be more easily developed than the pure delta. “chap08” — 2003/3/10 — page 211 — #10 Project study: advanced deep interdiction aircraft 211 Fig. 8.2 Design concept – conventional straight wing Fig. 8.3 Design concept – delta/diamond A decision matrix method was used to analyse the different options on a consistent basis. The criteria used to assess the options in the selection process are listed below together with (in brackets) the significance (weighting) to the overall assessment. Effectiveness of incorporating stealth technology into the layout (5) Aerodynamic efficiency (mainly L/D ratio) of the layout (5) Potential for low-weight design (4) Technical difficulties (ease of analysis) and risk (3) “chap08” — 2003/3/10 — page 212 — #11 212 Aircraft Design Projects + + + Fig. 8.4 Design concept – blended delta Field performance and rough ground handling (2) Maintainability and operational dependability (2) Survivability and ease of repair (2) Multi-role capability (1) Naturally, the choice of criteria and the relative weightings is highly subjective but a group response tends to smooth the assessment process. The result of the ‘voting’ on the criteria above is shown below: Conventional option (56), Delta/diamond layout (72), Blended body (58) The necessity for high L /D ratio and improved stealth were the key issues in theselection of the delta/diamond layout. It was also agreed that as much effort as possible should be given to the use of blending the profiling of the body (as on the B-2 aircraft). It was also decided that an advantage would be gained if the aircraft length was increased. These issues led to changes in the original configuration. To reduce aircraft maximum sectional area and effectively lengthen the aircraft, tandem seating and tandem weapon stowage was employed. This resulted in the concept sketch shown in Figure 8.5. The basic structural framework consists of a continuous (tip-to-tip) wing box. The weapons and main landing gear are suspended below this and housedin profiled fairings with radar reflective door and hinge edgings. Forward of the weapon bay, the profile is extended to accept the engine intakes which sweep up in S-bends to the top wing surface. This duct-profiling protects the intake profile against radar reflections from the engine compressor face. It also ensures clean airflow to the engines with the aircraft at high-incidence attitude. The nose landing gear is retracted into the space between the separate intake ducts. The twin engines are supported in cradles above the wing “chap08” — 2003/3/10 — page 213 — #12 Project study: advanced deep interdiction aircraft 213 Fig. 8.5 Selected and revised concept sketch structure. Nozzle exhaust ducts terminate forward of the wing trailing edge to shield the aircraft from downward infrared emissions. Fuel tankage is provided between the engine support cradles and intake ducting. The pilot and equipment bays are located in the aircraft centre line fuselage profile forward of the fuselage fuel tanks. The upper body is profiled to blend smoothly into the wing surface and to give an advantageous Sears–Haack volume distribution. 8.5 Initial sizing and layout The initial sizing of the preferred configuration requires estimates of the main aircraft parameters. Instead of just guessing these values it is agood idea to investigate thevalues “chap08” — 2003/3/10 — page 214 — #13 214 Aircraft Design Projects Table 8.1 Parameter Fighters Strikers Bombers Empty mass ratio (M E /M TO ) 0.45–0.60 0.41–0.54 0.37–0.42 Fuel mass ratio (M F /M TO ) 0.21–0.33 0.17–0.33 0.40–0.62 Payload ratio (M PAY /M TO ) 0.21–0.28 0.18–0.37 0.14–0.19 Wing loading (M TO /S) kg/sq. m 262–467 315–544 447–516 Thrust/Weight ratio (dry) 0.65–1.29 0.56–0.88 0.26–0.40 associated with existing aircraft of the same type. It is possible to compile a list of design data for existing military aircraft using published data. 6 The problem with using this approach for our project is the unique nature of the specified mission requirements of the design. It does not follow the ‘fighter’ class of aircraft because of our need to fly a longer range and carry a heavier weapon load than is normal for fighters. It does not fit into the ‘bomber’ class due to the higher speed and lower weapon load of our aircraft. ‘Multi-role’ and ‘strike’ aircraft may have some comparable features but these usually have much better manoeuvring ability and are not expected to supercruise for long periods. Using data on appropriate military aircraft from reference 6 (with extreme values ignored), it is possible to assess the variation of some design parameters (Table 8.1). It is clear from this analysis that there is wide variation in the aircraft used in the study. Also, as with all published data, the definition of aircraft parameters (e.g. empty weight) may notbe consistent from each manufacturer. The data therefore only provides a crude guide to the selection of parameters for use in the initial estimates. This implies that the initial estimates will be unreliable. It will be necessary to adopt a more refined analysis as quickly as possible. Some thoughts about our design that might help us to select suitable starting values: • Most of the aircraft in the survey are not supersonic in dry thrust so our design is likely to require a higher thrust to weight ratio than the bomber values. • Travelling for long distance at supersonic speed will require more fuel than is seen in the fighter and strike classes above but not as much as the max. bomber (B-52) value. • The fuel capacity required will be larger than on equivalent size aircraft so it may be advantageous to have a larger wing area to provide extra tankage. • A large wing size (low wing loading) will help in meeting the icy runway requirements. • The payload carried by our design, as defined in the specification, will give a relative low useful load ratio and the range flown at supersonic speed will give a high fuel mass ratio. • The empty mass ratio would also be reduced due to the large fuel mass but to account for the stealth requirement extra structure mass (radar absorbent materials) will be required. With no better information these two effects will be assumed to cancel each other, giving a conventional empty mass ratio. With these thoughts in mind, our initial estimates are shown below: • Empty mass ratio = 0.44 (this is low for fighters but high for bombers). • Fuel mass ratio = 0.46 (this is outside the range for fighter/striker aircraft but about average for bombers). “chap08” — 2003/3/10 — page 215 — #14 Project study: advanced deep interdiction aircraft 215 • Using the above assumption would make the payload ratio = 0.1 (as predicted, this falls below all aircraft classes). • Wing loading = 390 kg/sq. m (about 80 lb/sq. ft) (which is low for bombers, high for fighters and about average for strike aircraft). • Thrust loading = 0.60 (this is low for strike and fighter aircraft but it is not clear from the collected data how many of the sample have quoted afterburning (wet) thrust. It is outside the range for bomber aircraft). It is now possible to use the assumed values to make our first ‘rough’ predictions of the size of the aircraft: • From the problem specification we can predict that the payload (including two crew) is 6600 kg (14 550 lb). As we assume above that this represents 0.1M TO , the aircraft maximum take-off mass must be 66 000 kg (145 500 lb). • With an empty mass ratio of 0.44M TO , the empty mass = 29 000 kg (64 000 lb). • With a fuel mass ratio of 0.46 M TO , the fuel mass = 30 360 kg (67 000 lb). • With a wing loading of 390 kg/sq. m = 3826 N/sq. m (about 80 lb/sq. ft), the gross wing area = 170 sq. m (1827 sq. ft). • With a thrust loading of 0.6W TO , the total engine thrust (sea level, static, dry) = 388 kN (87 300 lb). This equates to 194 kN (43 700 lb) per engine. This makes our aircraft heavier and larger than any of the fighter and strike aircraft surveyed but much smaller than the existing bombers. The diamond planform (area, S = 170 sq. m, 1830 sq. ft) which is limited in span (b) to 18.3 m (60 ft) (to keep within the hangar width) will have a centre line chord = (2S/b) = 18.6 m. For a symmetrical planform (90 ◦ at the tip) the wing sweep is only about 45 ◦ and we must have at least 51.3 (see section 8.2.3). It is also advantageous to maintain a long overall length to reduce wave drag. Both of these requirements can be met by reducing aircraft span to 17 m (55.7 ft). In this case the centre line chord will be increased to 20 m (65.6 ft). Providing a 90 ◦ angle between the leading and trailing edges at the tip gives a 60 ◦ wing leading edge sweep angle. This geometry may need to be changed later in the design process if more fuel tankage is required. Using the concept sketch (Figure 8.5) and the values above we can now produce our first scale drawing of the aircraft (Figure 8.6). 8.6 Initial estimates With an accurate drawing of the aircraft (Figure 8.6) it is possible to estimate the component masses and drags (and lift). The predicted thrust will allow us to select a suitable engine or scale an existing design to provide engine performance data at all flight conditions. With mass, aerodynamic and propulsion data it will be possible to perform initial performance calculations and draw our first constraint diagram. “chap08” — 2003/3/10 — page 216 — #15 216 Aircraft Design Projects Fuel O/length = 20 m O/height = 4.8 m Span=17m Area (ref.) = 170 m 2 LE = 60° Fuel Fuel Eng. Eng. Equip. Cockpit Eng. Cockpit Intakes Eng. 10 m 30 ft Thrust vector Fig. 8.6 Initial baseline aircraft general arrangement 8.6.1 Initial mass estimations The initial mass estimates can be calculated by using published empirical equations based on existing aircraft designs. 4 As our aircraft has a unique operating envelope, such methods may be regarded as crude. At this stage in the design process, the analysis is likely to be more accurate than the ‘guesstimates’ made from the survey used above. Using our knowledge of the aircraft specification, some corrections to the method can be applied. All the required input data for the method can be gleaned from the initial layout drawing, the project specification and common sense. Applying such data to the equations in reference 4 gives the mass statement shown in Table 8.2. This initial estimate of MTO is substantially less than previously predicted. The main reason for this reduction is due to the lower prediction of aircraft empty mass. Although, as expected, the propulsion system mass is large due to the high thrust “chap08” — 2003/3/10 — page 217 — #16 Project study: advanced deep interdiction aircraft 217 Table 8.2 Mass Component (kg) (lb) % MTO Wing (inc. controls) 1 888 (4 163) 3.0 Body (inc. engine cowls) 7 070 (15 589) 11.3 Undercarriage (all units) 1 138 (2509) 1.8 Total structure 10 096 (22 262) 16.1 Propulsion system 12 077 (26 630) 19.2 Fixed systems 3 673 (8 100) 5.8 Aircraft empty 25 846 (56 990) 41.1 Crew (two pilots) 500 (1 100) Weapons 6 100 (13 450) Zero fuel mass 32 446 (71 543) 51.7 Fuel ∗ 30 360 (66 944) 48.3 Max. take-off mass 62 806 (138 487) 100.0 ∗ The fuel load is retained at the value estimated from the higher MTO mass originally predicted. This will need to be checked when the mission analysis is completed to weight ratio, the aircraft structure and fixed systems masses are low. This could have been expected as the compact and stiff structure framework will provide a light structure. However, for our high-tech, modern weapon system, the low systems mass must be treated as suspicious. As the project develops, and more detail is known about the aircraft systems, it will be necessary to reassess this estimate. As the aircraft empty mass estimation was based mainly on the original value of MTO it is expected that the aircraft mass and size could be reduced. Before any changes are contemplated, it is advisable to continue with the aerodynamic and performance estimations using the original design. In this way, all the design modifications can be assessed at the end of the initial estimation process. 8.6.2 Initial aerodynamic estimations The initial aerodynamic estimations concern the prediction of aircraft drag and lift. For this aircraft the main focus of drag will be on the supersonic wave drag (C Dw ) estimation. Using the wave drag equation in reference 4, with the following input values, gives the first estimation of C Dw : Aircraft cruise Mach number, M = 1.6 Aircraft max. cross-sectional area, (A max ) = 10.06 sq. m Reference wing area, S ref = 170 sq. m Wing LE sweep = 60 ◦ Aircraft overall length (less any constant sections), L = 20 m An adjustment factor to relate the actual cross-section distribution to the Sears– Haack perfect shape, E WD = 1.4 (assuming a smooth distribution from the blended body) Gives , C Dw = 0.02104 This is a very large drag increment that will substantially penalise the design. Somehow, we will need to either reduce the cross-sectional area or increase the aircraft length. The area cannot be changed significantly unless we alter the internal requirements. It [...]... #32 233 234 Aircraft Design Projects Table 8.5 Component lb kg 6 1 57 832 5 210 2 835 868 2 633 2 800 378 2 368 1 289 394 1 1 97 18 535 8 406 16.2 15 600 76 9 1 054 7 091 349 478 13 .7 0 .7 0.9 PROPULSION 17 423 8 72 8 15.3 18.8 Fuel system and tanks Aircraft systems Avionics Cockpit systems Weapon systems 1 72 3 1 546 2 370 1 440 1 500 78 3 70 1 1 077 653 682 1.5 1.4 2.1 1.3 1.3 19.5 14.0 9.5 10.0 17. 0 Wing Control... 228 — # 27 Project study: advanced deep interdiction aircraft Table 8.3 TE sweep (forward) LE sweep 65 70 25 15.18∗ 17. 53 13.98 19.03 12.66 21.03 60 30 20 15.56 17. 10 14. 27 18.63 12. 87 20. 67 15.93 16 .70 14.56 18. 27 13.08∗ 20.34 30° LE sweep 60° 17. 53 m 15.18 m 15 10 16.31 16.31 14.85 17. 91 13.28 20.03 16 .70 15.93 15.14 17. 57 13.49 19 .72 10° 20° Centre line chord 16 .70 m span 15.93 15.93 m 16 .70 30°... (SSL) t /s 0.8 Approach speed F-22 uv F-23 Original design point 10 150 0 f ft /s t /s Revised design point 0 .7 0.6 F-32 0.5 al TO F-1 17 Norm 0.4 F-14DX Dash + Climb F-16XL 0.3 2000 2500 3000 B-2 3500 4000 Concorde 4500 5000 Supercruise 5500 6000 Wing loading (N/sq m) Fig 8.10 Constraint diagram “chap08” — 2003/3/10 — page 223 — #22 F-15EX 6500 70 00 223 224 Aircraft Design Projects (a) For the approach... % MTO 5.4 0 .7 4.6 3.2 (u/c) 2.3 Fuel Max take-off 7. 6 1 100 13 448 500 6 113 1.0 11.8 59 082 27 623 51.9 25 000 48.1 51 73 9 18.0 23.5 16.5 19 .7 6 .7 13.5 39.1 114 082 ZERO FUEL 3 891 21 025 55 000 Crew and op items Weapons 8 579 44 5 37 EMPTY Arm (m) 100.0 10.0 17. 5 18.0 (central) 20.5 (wing) estimated positions of the component masses With the aid of a spreadsheet, various combinations of aircraft loading... 6 Intake 20 Intake E-E F-F D-D C-C Fuel A-A B-B B C D E A F A/c CG A B C D E F 17 Fig 8.13 Revised baseline GA provide a useful storage volume to house sensitive sensors and flight instruments The increased centre line chord provides sufficient volume to accommodate the weapon bay in a four-across configuration and part of the engine depth This produces a blended body shape for the aircraft For stealth... V = aircraft speed at the condition under investigation h = aircraft altitude at the case under investigation ρ = air density at height h CDO = aircraft zero-lift drag coefficient k1 = aircraft- induced drag coefficient n = aircraft normal load factor = L/W dh/dt = aircraft rate of climb at the case under investigation g = standard gravitational acceleration = 32.2 ft/s2 (or 9.81 m/s2 ) dV /dt = aircraft. .. MTOM of 60 806 kg (134 077 lb) This makes the new wing area = (60 806 × 9.81)/4500 = 133 sq m approx (i.e 1425 sq ft) The static sea-level military thrust (both engines) = (60 806 × 9.81) × 0.58 = 346 kN (77 78 2 lb) SSL thrust per engine = 173 kN (38 900 lb) It is now possible to modify the original aircraft general arrangement drawing and to make some detailed estimates for the aircraft mass, aerodynamic... separation of the airflow at the nose of the aircraft, it will add length which will reduce wave drag, and it will “chap08” — 2003/3/10 — page 230 — #29 Project study: advanced deep interdiction aircraft Table 8.4 Weapon name Guidance Number Size, length × dia (m/ft) Configuration GBU- 27 Mk-84 LDGP JDAM AGM-154 JSOW SSB Laser None GPS Internal N/A 4 4 4 4 16 4 .7/ 15.4 × 0 .76 /2.5 3.6/11.8 × 0.45/1.5 3.0/10.0 ×... most significant of these are the take-off and landing phases In addition to the aircraft clean condition we must add landing gear, flap and any aerodynamic Sears–Haack ideal distrubution Canopy and cockpit Rear engine installation Nose Tail Aircraft length Fig 8 .7 Sears–Haack cross-sectional area distribution “chap08” — 2003/3/10 — page 219 — #18 219 220 Aircraft Design Projects retarding devices (e.g lift...218 Aircraft Design Projects is relatively easier to increase the length (see later aircraft drawings) Assuming that it is possible to stretch the aircraft to 28 m (92 ft) the calculation above would change to: C Dw = 0.01408 The parasitic drag will be estimated by using an equivalent skin friction coefficient of 0.0025 (representative of a smooth fast transport aircraft) Hence, with an estimated aircraft . two AIM-120. • Four GBU- 27 + two AIM-120. • Four 2000 lb JDAM + two AIM-120. • Four AGM-154 JSOW + two AIM-120. • Sixteen 250 lb small smart bombs. “chap08” — 2003/3/10 — page 210 — #9 210 Aircraft. 8.10. 0.3 0.4 0.5 0.6 0 .7 0.8 Thrust loading (SSL) 2000 2500 3000 3500 4000 4500 5000 5500 6000 6500 70 0 0 Wing loading (N/sq. m) 0.9 1 1.1 F-1 17 F-16XL F-15EX F-14DX Normal landing Approach speed Original design point Revised. involved in detail design at this early stage in the design of the aircraft. “chap08” — 2003/3/10 — page 226 — #25 226 Aircraft Design Projects (d) Normal take-off For take-off conditions the