Aircraft Design Projects Episode 11 pdf

“chap09” — 2003/3/10 — page 281 — #12 High-altitude, long-endurance (HALE) uninhabited aerial surveillance vehicle (UASV) 281 Fig. 9.7 Flying-wing layout option production aircraft (with the exception of the B-2 stealth bomber). Many enthusiasts for the type have claimed the layout to be aerodynamic and structurally efficient but it seems that such expectations have not yet been realised (the B-2 layout is selected for stealth reasons). The reason may be due to the linking of stable torsional deflections to the aerodynamic forces and the consequential requirement to modify the wing planform and sectional geometry to avoid this problem. The layout is regarded as efficient at one design point but seriously compromised away from this condition. The flying wing con- figuration was considered in the German study 1 but dismissed on these technical issues. 9.6.4 Braced wing layout (Figure 9.8) The structural bracing of the wing to the fuselage was a common feature in historic aircraft layouts. This was done to reduce the loads in the wing to match the relatively poor structural properties of the materials used in the construction. The development of stronger and more consistent materials allowed such bracing and the associated drag penalty to be eliminated. The traditional monoplane wing layout has been the preferred choice over the past several decades. As wing aspect ratio is increased, the benefit of bracing becomes more attractive as it significantly reduces wing bending moments. In recent years, some NACA funded research 6 has shown that wing bracing could provide advantages to the design of long-range civil jet transport aircraft. The purely tensile loaded brace reduces the shear, bending and torsion on the wing structure. This correspondingly allows either a thinner wing or a larger aspect ratio to be used on the wing geometry. A thinner wing would allow the wing to be less swept for a design critical Mach number. All of these effects reduce aircraft drag and consequently fuel burn. Positioning the brace attachment to the wing ahead of the sectional structural axis also provides a reacting nose-down moment to stabilise the divergence tendency associated with a swept forward wing planform. Mounting the wing on the fin structure and adding dihedral to counter the unstable yaw coupling from the swept forward planform places the wing well above the ground “chap09” — 2003/3/10 — page 282 — #13 282 Aircraft Design Projects Fig. 9.8 Braced-wing layout option plane during landing. This provides the aircraft with adequate bank angle to protect against disturbed landing manoeuvres. The main drawback with the configuration is associated with the novelty of the layout and its potential for technical risk. 9.6.5 Configuration selection Of the four options, only the conventional and braced wing seems to be worth further consideration. As the German design study 1 selected a conventional layout for their baseline design we will investigate the braced wing layout. This will provide a useful comparison with the previous study. Having selected the braced wing layout there are several detail design considerations to be made: 1. The engine mountings, fuselage brace attachments and the main undercarriage mounting will be combined into a central fuselage structural framework. This will leave the forward fuselage structure uncluttered and capable of holding the equipment modules as conformal containers below the fuselage structural beam. 2. To avoid the difficulty of attaching the brace to the wing structure, and the possibility of complex airflows at the junction, pylon mounted equipment/fuel pods will be installed on the wing. The brace will be attached to the pod support structure (see Figure 9.9). 3. The brace structure will need to be streamlined and this will provide the opportunity to run equipment service lines or fuel supply pipes directly between the wing and fuselage. 4. It may be possible to use the wing and brace structures to house conformal radar antenna (as proposed by Boeing on their CSA). 5. To reduce trim drag (an important feature on long-range/endurance aircraft) the forward fuselage could support a small canard surface mounted above the equip- ment modules. Care will need to be taken on the position of this surface relative “chap09” — 2003/3/10 — page 283 — #14 High-altitude, long-endurance (HALE) uninhabited aerial surveillance vehicle (UASV) 283 A A (a) Connection without attachment pod. Flow interference at win gj oint A. (b) Connection using attachment pod. Improved flow at j unction A. Fig. 9.9 Wing to brace interconnection detail to the engine intakes. Wind tunnel tests will need to be done to finalise the exact geometry. 6. It will be necessary to incorporate wing inboard control surfaces to provide pitch control. 7. Although main wheels will be required, it may be preferable to use skids for the third (nose or tail) unit. 9.7 Initial sizing and layout While the aircraft is of an unconventional configuration, the initial sizing and lay- out process will follow the normal procedure. This will involve estimating the aircraft take-off mass, wing loading, some airspeed predictions, wing layout and powerplant sizes. These are described in the following subsections. Finally, all of the component studies are linked together to produce the initial baseline aircraft layout. 9.7.1 Aircraft mass estimation In order to size the aircraft it is necessary to estimate the maximum take-off mass. The formula below is often used for this purpose (see Chapter 2, section 2.5.1 for the definition of terms): M TO = (M UL )/{1 − (M E /M TO ) − (M F /M TO )} For the HALE aircraft there are some difficulties that arise from the definitions of aircraft systems to be included in the aircraft empty mass ratio. Many of the systems on the aircraft are directly related to the type of operation. Some of the equipment may be changed to suit the mission (reconnaissance, communication, surveillance, atmospheric research and monitoring). To resolve the lack of knowledge of these systems and the variability with the mission, the equipment mass will be assumed to be 800 kg. This value will be attributed to the ‘useful load’ in the above equation. At a later stage in the development of the design it will be appropriate to conduct sensitivity analyses around this assumption. A second difficulty arises due to the expected, unusually large, fuel ratio. An aircraft with a duration of 24 hours is almost unique. Therefore data from other, shorter- range aircraft may be misleading. For this reason it will be essential to check the fuel requirements as soon as the aircraft mass, lift, drag and engine characteristics are known with reasonable accuracy. “chap09” — 2003/3/10 — page 284 — #15 284 Aircraft Design Projects The main conclusion from these observations is that the estimation of aircraft max- imum mass from the above expression must be treated with suspicion and regarded as tentative. Several iterations of the subsequent analysis will be necessary before confidence in the results can be realised. Analysis of the technical descriptions of the main aircraft types described in section 9.5 shows the following values for empty mass ratios: U-2S 0.445 Stratos 1 0.570 Stratos 2 0.500 Stemme 0.680 These contrast with a value of 0.254 for the German EADS project. 1 Our aircraft design brief is closer to the U-2S and EADS aircraft. The U-2S is recognised as being an old aircraft design with 1950/1960’s materials and construction methods. Application of modern materials and methods would be expected to reduce the structural mass. The EADS aircraft data may have linked more of the equipment mass to the useful load component than assumed in our case. Without more information, it is difficult to choose between these two values for empty mass ratio, therefore an average figure of 38 per cent will be initially used for our design. The fuel fraction can be estimated using the Breguet range equation: (M F /M TO ) = (engine cruise sfc) ·[1/(L/D)]·(flight time) Note: engine sfc varies with cruise altitude. For a typical medium bypass ratio turbofan engine, the following relationship is quoted 7 : (sfc) altitude /(sfc) sea level = θ 0.616 where θ is the ambient air temperature ratio (T A /T SL ). In the stratosphere the ISA temperature is constant at 216.76 K. ISA sea-level temperature is 288.16 K. This makes θ = 0.75. Hence, (sfc) altitude = 0.84(sfc) sea level A medium BPR engine is likely to have a sea-level sfc of 0.55 (lb/lb/hr or N/N/hr). Therefore using the above formulae gives an engine sfc in the stratosphere of 0.46. With a high aspect ratio wing and slender fuselage the aircraft lift to drag ratio (L/D ) in cruise could safely be assumed to be better than the value of 17 which is typical of modern civil airliners. Due to the forward swept wing and the interference arising from the brace structure it will not be possible to achieve the value of 40 which is typical of high performance gliders. Being conservative, we will assume a value of 25 but this will need to be checked and adjusted when detailed drag estimations are available later in the design process. The duration of the patrol is specified as 24 hours in the design brief. It is unclear if this is to include the time needed to reach the patrol area, so an extra two hours will be added to this time. A design duration of 26 hours will be used in the analysis below: Hence, (M F /M TO ) = (0.46) · (1/25) · (26) = 0.48 We will add 10 per cent for contingencies to give a design value of 0.53. Using the above values in the initial aircraft take-off mass equation gives: M TO = 800/(1 − 0.38 − 0.53) = 8888 kg (19 600 lb) To provide for some design flexibility in the subsequent work a design (max.) mass of 9200 kg (20 280 lb) will be assumed. “chap09” — 2003/3/10 — page 285 — #16 High-altitude, long-endurance (HALE) uninhabited aerial surveillance vehicle (UASV) 285 9.7.2 Fuel volume assessment The calculation above predicts a fuel mass of 0.53 × 8888 = 4693 kg (10 350 lb). With a specific mass for aviation fuel of 0.8 (fuel varies between 0.76 and 0.82), this mass will need 5833 litres (5.866 m 3 , 207 ft 3 ) tankage volume to hold the fuel. When the wing geometry has been defined, a check will be necessary to establish if the volume can be accommodated in the integral wing tanks. If not, the size of other storage tanks to be included in the aircraft layout will need to be determined. 9.7.3 Wing loading analysis Flying at high altitude where the air is thin requires either a fast airspeed or a high value of lift coefficient (or probably a combination of both) to reduce the required wing area. Maximising these parameters for a chosen altitude sets the value for the maximum wing loading as shown below: L = 0.5ρV 2 SC L Substituting ρ = p/RT , and using a = (γ RT ) 0.5 , where (ρ) is air density, ( p)isair pressure and (a) is the speed of sound. Assuming γ = 1.4 and using M = Mach number (=V /a), gives the equation: L/S = 0.7pM 2 C L Typical values of the parameter (M 2 C L ) max , range from 0.05 for gliders to 0.6 for mili- tary jets. Conventional civil transports lie in the range 0.2 to 0.4 (i.e. M0.8 @ C L = 0.4 gives (M 2 C L ) = 0.24). Figure 9.10 shows the distribution of maximum wing loading against altitude for various values of (M 2 C L ) max . The areas marked for each type rep- resent the common values. Obviously some aircraft are designed to operate away from these regions. Figure 9.11 shows the portion of the previous figure relating to high altitude operations. As discussed in section 9.3, calmer wind conditions are found at 0 0 5 10 15 20 2000 4000 6000 8000 10 000 12 000 Cruise altitude (km) Wing loading (N/m 2 ) 0.2 0.3 0.4 (Ma 2 C L ) max Gliders Civil airlines Military jets UASVs Fig. 9.10 Max. wing loading versus altitude “chap09” — 2003/3/10 — page 286 — #17 286 Aircraft Design Projects 0 15 16 17 18 19 20 21 22 500 1000 1500 2000 2500 3000 Cruise altitude (km) Wing loading (N/m 2 ) 0.2 0.3 0.4 (Ma 2 C L ) m Project design point UASV operating range . . Fig. 9.11 UASV wing loading selection altitudes around 18 km (59 000 ft). This sets the design point shown in Figure 9.11. Therefore, the selected wing loading is 1800 N /m 2 (183.5 kg/sq. m, 37.6 lb/sq. ft).For our chosen aircraft design mass of 9200 kg (20 280 lb), this equates to a minimum wing area of 50 m 2 (537 sq. ft). 9.7.4 Aircraft speed considerations The aircraft operating envelope is bounded at slow speed by the aircraft stall perfor- mance. At high speed, the operating envelope is restricted by the available engine thrust, the rise in transonic wave drag and the effects of the associated buffet on the aircraft structure. The effect of high altitude operation affects both of these speed boundaries. For a given wing area and sectional C Lmax value, the stall speed will increase as air density reduces as defined below: V stall =[L/(0.5ρSC Lmax )] 0.5 As air temperature reduces with altitude (up to the start of the stratosphere) the speed of sound and thereby the aircraft speed at the onset of transonic flow will reduce. The speed of sound is determined by the relationship: a = a o θ 0.5 where (a o ) is the speed of sound at sea level = 340.29 m/s, 661 kt. These effects are shown diagrammatically in Figure 9.12. It is advisable to fly at a speed greater than the stall speed to allow a margin of safety to protect against gusts. This margin will avoid inadvertent stalling and reduce pilot/system control demands. This defines the minimum speed boundary. For many types of aircraft the margin is set by applying a factor of 1.3 to the stall speed in the low speed, approach to landing, phase. High wind speeds may demand an increase in this margin. Although wind speed does not affect the aerodynamic parameters of the aircraft (the aircraft travels with the ambient air and only relative changes are significant) it does alter the perceived (relative to the ground) climb and descent flight “chap09” — 2003/3/10 — page 287 — #18 High-altitude, long-endurance (HALE) uninhabited aerial surveillance vehicle (UASV) 287 Feasible flight regime Flight safety zone 1.3 V s Aircraft stall (V s ) boundary Aircraft s p eed (m/s TAS) 24 20 16 12 8 4 0 100 200 300 400 Altitude (km) Troposphere Stratospher e Speed of sound M1.0 boundary Transonic flow zone: buffet boundary Fig. 9.12 Operating speeds constraints (diagrammatic) 10 50 100 150 200 250 12 14 16 18 20 22 24 Aircraft speed (m/s TAS) Altitude (km) Selected aircraft design/ operational point (H =18km V =210 m/s) Transonic flow and buffet Flight safety speed boundary = 1.3 V s Aircraft stall boundary (V s ) (M =0.9M TO , C L M = 1.4) Feasible operating region 80% wind vector added to V s boundary Fig. 9.13 Aircraft speed envelope paths. This will influence the time needed to get onto, and from, the operating station. The aircraft high-speed boundary is directly affected by the aerodynamic (transonic) characteristics of the aircraft (mainly the wing geometry and pressure interference effects). Smooth aircraft cross-section area shaping, supercritical wing profiles and increased sweep are methods to delay the onset of transonic effects. Civil transports push the boundary to about M0.85 but this increases drag by about 3 to 5 per cent. Reducing the operating speed to less than M0.8 should avoid this penalty. Figure 9.13 shows the absolute (stall and Mach1.0) boundaries for the aircraft together with the 80 per cent average wind speeds at various heights. The wind speed at altitude is important as it will add or subtract to the ground speed and therefore the “chap09” — 2003/3/10 — page 288 — #19 288 Aircraft Design Projects search pattern. Note that at about 21 km, the minimum and maximum boundaries are nearly coincident. To fly above this height would require the aircraft to reduce weight, have an increased wing area or an increase in C Lmax , or combinations of these changes. For a given aircraft geometry, a cruise-climb technique, in which height is gained as the aircraft mass reduces with fuel use, could be considered. The design point selected on the above considerations and the earlier discussion (section 9.3) is: Operating altitude (initial) = 18 km (59 000 ft) Operating speed = 210 m/s (408 kt), representing M0.71 at 18 km At the above condition the aircraft lift coefficient is 0.604 at the start of patrol (mass = 0.9M TO ). This reduces to 0.332 at the end of the patrol (mass = 1.3M empty ). At the end of patrol, the aircraft stall speed will have reduced from an initial value of 138 m/s (268 kt) to 102 m/s. This change would allow the aircraft to fly progressively either slower or higher. The discussion above has concentrated on the cruise perfor mance; it is also neces- sary to check the approach speed to determine if it is acceptable. Assuming ISA-SL conditions with an aircraft mass on approach of 1.15M empty and a C Lmax of 1.4 (i.e. no flaps): V 2 s =[1.15 (0.38 · 9200) 9.81]/[0.5 · 1.225 · 50 · 1.4] V s = 30.3 m/s (59 kt) If V approach = 1.3V stall , then V approach = 1.3 × 30.3 = 39.4 m/s (75.6 kt) This speed should be slow enough to allow for automatic/remote landing control in the UAV version. For emergency landing at higher weight, consideration may need to be given to the provision of fast fuel dumping. 9.7.5 Wing planform geometry Selection of the wing planform is the most significant design decision with regard to the aircraft performance. This aircraft will spend most of its time on long-duration patrol missions. It is therefore important to choose the wing geometry to ‘optimise’ this part of the operating envelope. In this case, drag reduction forms the main basis for the selection of wing characteristics. In the search phase, the aircraft induced drag will form a significant component of drag. Selecting a high aspect ratio for the wing is an effective method of reducing induced drag. On conventional monoplane designs, high values for aspect ratio lead to a substantial increase in wing structural mass. This penalty arises due to the outboard movement of the centre of lift (away from the wing root/bodyside attachment). The bracing structure selected on our design avoids this penalty as part of the outboard lift is reacted by the brace structure. Higher values of wing aspect ratio than normally seen on conventional designs are therefore feasible. For a given wing area, high aspect ratio corresponds to a large span, it may there- fore be necessary to impose a limit to ensure that the aircraft is easy to handle on or near the ground. In this respect, an aspect ratio of 25 will be selected. This com- pares to values in the range 7–9 for civil transports and 20–30 for higher performance gliders. “chap09” — 2003/3/10 — page 289 — #20 High-altitude, long-endurance (HALE) uninhabited aerial surveillance vehicle (UASV) 289 Airload (a) Traditional monoplane (cantiliver) Max. BM Max. bending moment diagram Note the wing progressively reacts an increasing bending moment from tip to root chord. Airload Brace load Max. bending moment diagram Note the significant reduction in maximum bending moment reacted by the wing structure. (b) Braced wing structure Max. BM Fig. 9.14 Wing bending-moment diagrams To delay the onset of transition the wing will be swept forward by 30 ◦ . A thin wing (8 per cent) thickness will be adopted. These characteristics should provide a criti- cal Mach number above M0.84 and therefore avoid transonic wave drag penalty and structural buffeting in the cruise/search phases. The high aspect ratio wing will produce a small chord and correspondingly a low value for the airflow Reynolds number. This will encourage the retention of laminar flow over the wing section. A transition at 70 per cent chord may be possible if the wing profile skins are smooth, continuous (no gaps or junctions) and the surfaces are kept clean. This should be possible with a composite construction and normal military service care. The reaction force on the wing from the brace will alter the wing bending moment distribution. This will cause an unusual distribution of wing taper. In a traditional unbraced design the maximum bending moment occurs at the wing to fuselage attach- ment section (see Figure 9.14). This is the position where the deepest wing thickness is required and therefore the widest chord. A straight tapered wing planform with the largest chord at the root is the usual configuration. For the braced wing layout in which the relative stiffness of the wing structure and the brace can be selected, the largest wing bending moment may be at the brace attachment section. To reflect this change of bending moment distribution the wing taper will be unconventional. The largest chord will be at the brace attachment position, as shown in the initial aircraft layout drawing. The aircraft wing geometry is now defined: Wing area 50 sq. m 537 sq. ft Wing aspect ratio 25 Wing span (=(S · A) 0.5 ) 35.4 m 116 ft Wing sweep 30 ◦ forward Physical span (=35.4 cos 30) 30.6 m 100 ft Wing thickness ratio 0.08 Wing max. C L 1.4 Wing taper to match brace geometry Wing section supercritical cambered “chap09” — 2003/3/10 — page 290 — #21 290 Aircraft Design Projects 9.7.6 Engine sizing As already described, the engine type will be a medium BPR turbofan. The required thrust will be dependent on the climb performance at the cruise altitude. If the climb rate at high altitude is too small then the time required to reach the search height will be too long. Without a specified value for climb rate, we will assume the civil transport criterion of 300 ft/min (1.53 m/s). This assumption will need to be checked when more detailed mass, aerodynamic, propulsion and performance analysis is possible. A sensitivity analysis will be required to show the inter-relationship of this assumption with the commonly specified 100 ft/min service ceiling criterion. The fundamental analysis is shown below: From basic flight mechanics: (dh/dt) = (T − D) · V /W (9.1) To obtain the required sea level, take-off thrust/weight ratio, equation (9.1) is modified to: (T /W ) SL = (β/α)[(D/W o ) + (1/V )(dh/dt)] A (9.2) where subscripts SL is sea level, o is initial (take-off) value and A refers to altitude values. (β) represents the mass fraction at the start of cruise (which, in this case we will assume to be 0.85). (α) represents the engine thrust reduction with altitude (T A /T SL ). From Eshelby’s book 8 :(T A /T SL ) = σ 0.7 in the troposphere (up to 11.02 km) and σ 1.0 in the stratosphere (σ is relative density = (ρ altitude /ρ SL )). This assumes a constant engine rating. For this type of aircraft the loss of thrust at this high altitude will be large therefore it is likely that the static sea thrust (T SL ) at the climb (or cruise) rating will be suitable to meet the take-off requirement. For this reason, a constant (climb or cruise) rating will be assumed. As our cruise will be in the stratosphere: (T A /T SL ) = (ρ 11.02 /ρ SL ) 0.7 · (ρ A /ρ 11.02 ) = 0.24 The values for ρ (air density) can be found in ISA tables (in SI units, ρ SL = 1.225, ρ 11.02 = 0.364 kg/cu. m) (in Imp. units, = 0.002378, 0.000707 slug/cu. ft) Aircraft drag at the cruise condition = 0.5ρ A V 2 SC D where V = 210 m/s (408 kt), S = 50 sq. m (537 sq. ft), and C D assumed to be 0.022. Aircraft take-off weight W o = 9200 × 9.81 = 90.25 kN = 20 280 lb Equation (9.2) is computed and plotted in Figure 9.15. At our selected design altitude of 18 km (see Figure 9.11) the required thrust to weight ratio (known as the thrust loading) is 0.24 for a 300 ft/min climb ability. The corresponding line at 100 ft/min indicates a service ceiling, with this thrust ratio, of 24 km (78 700 ft). From our previous discussions, this value appears to be satisfactory. This thrust loading gives a required take-off thrust of: T o = 0.24 · 9200 · 9.81 = 21.6 kN = 4870 lb With two engines this equates to 10.8 kN (2434 lb) per engine. It is necessary to check that this thrust is adequate for safe single-engine take-off in an emergency. The civil aircraft airworthiness requirement sets a climb gradient of 0.024 at 50 ft height and speed V 2 (undercarriage retracted but take-off flaps still deployed). [...]... al., Civil Jet Aircraft Design, Butterworth-Heinemann, 2000, ISBN 0-340741-52-X 8 Eshelby, M E., Aircraft Performance, Theory and Practice, Butterworth-Heinemann, 2001, ISBN 0-340758-97-X 9 Howe, D., Aircraft Conceptual Design Synthesis, Professional Engineering Publishing, UK, ISBN 1-86058-301-6 10 Nicholai, L M., Fundamentals of Aircraft Design, METS Inc., San Jose, California 95120 11 Rabinowitz,... 293 — #24 293 294 Aircraft Design Projects 9.8 Initial estimates With a fully dimensioned general arrangement drawing of the aircraft available it is possible to undertake a more detailed analysis of the aircraft parameters This will include component mass predictions, aircraft balance, drag and lift estimations in various operational conditions, engine performance estimations and aircraft performance... gear Total structure Propulsion group Fixed equipment Aircraft empty Useful load Fuel load Aircraft MTO lb 1082 277 883 147 409 2798 820 736 4354 800 4695 9849 2 386 611 1 947 324 902 6 170 1 808 1 622 9 600 1764 10 352 21 716 % MTO 11. 0 2.8 9.0 1.5 4.1 28.4 8.3 7.5 44.2 8.1 47.7 100.0 “chap09” — 2003/3/10 — page 297 — #28 297 298 Aircraft Design Projects Datum (plan) Datum (side) Forward fixed equipmt... m/s (107 kt) FN = 11. 37 kN (2556 lb) D = (0.5ρV 2 ) SCD = 1853 × 50 × 0.03819 = 3538 N (795 lb) RoC = [55.1/(9200 · 9.81)] (113 70 − 3538) = 4.78 m/s (940 fpm) “chap09” — 2003/3/10 — page 301 — #32 301 302 Aircraft Design Projects We can also calculate the aircraft climb gradient = sin−1 (RoC/V ) = 0.08 (i.e 8 per cent which is satisfactory) (note: the minimum value for civil transport aircraft is 2.4... — 2003/3/10 — page 306 — #37 High-altitude, long-endurance (HALE) uninhabited aerial surveillance vehicle (UASV) 9 .11 9 .11. 1 Aircraft specification Aircraft description Aircraft type: Manned or uninhabited high-altitude, long-endurance, reconnaissance vehicle Design features: The novel aircraft layout, with a high aspect ratio, multi-tapered, swept-forward braced wing planform, provides a platform for... By-pass ratio = 3.3 “chap09” — 2003/3/10 — page 291 — #22 291 292 Aircraft Design Projects This engine will give about 20 per cent extra thrust than required for aircraft performance so should be adequate to meet the aircraft service needs 9.7.7 Initial aircraft layout The previous sections have set out the geometrical requirements for the aircraft It is now possible to produce the first general arrangement... calculated as T = 11. 37 kN (2556 lb) And specific fuel consumption (C): C/Co = [1 − (0.15 · 3.3)0.65 ][1 + 0.28(1 + 0.063 · 3.32 )0.7]σ 0.08 Giving, C = 0.493 9.8.6 Aircraft performance estimations Initial estimates of aircraft performance are based on methods described in most aircraft design textbooks (e.g references 7 to 10) Point estimates are required to determine the suitability of the aircraft layout... H., ‘Review of unconventional aircraft design concepts’, Journal of Aircraft 25, 5: 385–392 6 Ko, A et al ‘Effects of constraints in multi-disciplinary design of a commercial transport with strut-braced wings’ AIAA/SAE World Aviation Congress 2000/1, paper 5609 See also Gundlach, J T et al ‘Concept design studies of a strut-braced wing, transonic transport’ AIAA Journal of Aircraft, Vol 137, No 6, Nov-Dec... 7.2 9.7 119 .4 — 2.0 129.7 237.1 0.592 0.974 0.022 0.022 206.9 445.8 28.6 0.376 23.4 (with no height gain) = one engine inoperative at the start of climb, i.e emergency take-off case “chap09” — 2003/3/10 — page 299 — #30 299 300 Aircraft Design Projects At the initial cruise speed and height, the design lift coefficient will be 0.59, as shown in Table 9.3 Much more work would need to be done in designing... m (537 sq ft) for our aircraft Formulae used for the above estimation can be found in most aerodynamic or aircraft design textbooks (e.g reference 7) Geometrical inputs are scaled from the layout drawing The results (with a reference area of 50 m2 /537 sq ft) are shown in Table 9.3 9.8.4 Aircraft lift estimations To reduce complexity and to avoid drag increases in cruise, the aircraft will be manufactured . #23 292 Aircraft Design Projects This engine will give about 20 per cent extra thrust than required for aircraft performance so should be adequate to meet the aircraft service needs. 9.7.7 Initial aircraft. 0.254 for the German EADS project. 1 Our aircraft design brief is closer to the U-2S and EADS aircraft. The U-2S is recognised as being an old aircraft design with 1950/1960’s materials and construction. #25 294 Aircraft Design Projects 9.8 Initial estimates With a fully dimensioned general arrangement drawing of the aircraft available it is possible to undertake a more detailed analysis of the aircraft