in the break-up of an aircraft. It cannot be easily suppressed by a pilot control input, as oscillations typically occur at frequencies above 1 Hz (Nickel and Wohlfart (1994)).
For flutter to arise during subsonic flight there needs to be a coupling between at least two degrees of freedom, for example, spar-bending and wing torsion. Both tailed and tailless aircraft have many degrees of freedom, which means that both types are prone to flutter.
However, with the absence of a tailplane, Nickel and Wohlfart (1994) comment that tailless aircraft may actually be slightly more flutter resistant. In fact, in the early FW aircraft flight tests by Northrop, Northrop (1957) does not comment on any aeroelastically induced problems. However, Weishaar and Ashley (1973) suggest that the main reason for this may be due to the conservative structural design of the early FW’s.
Weishaar and Ashley (1973) carried out elementary aeroelastic studies to compare the behaviour of FW arrangements relative to conventional aircraft. In the idealised analysis, the FW was treated as a free-free beam-rod, with the conventional aircraft modelled as an equally stiff and massive cantilever beam-rod of half the span. On comparing these two simple models, in the absence of bending and torsion coupling, the analysis showed equal low torsional frequencies and pairs of higher harmonics. The analysis therefore suggests that it is unlikely for any new aeroelastic phenomena to present themselves with the FW layout.
Moreover, the overall conclusion after assessing the limitations of the various models is that flutter effects for FW aircraft seem more favourable.
As the lift and inertial distributions along the span of a tailless aircraft are approximately proportional to one another, Weishaar and Ashley (1973) expect the contribution to the overall internal structural load from payload and fuel-load, e.g. during a pull-up manoeuvre or gust encounter, to be mild. Therefore, on consideration of stiffness and aeroelasticity effects, Weishaar and Ashley (1973) expect that a higher structural weight fraction will result in order to attain acceptable flutter margins, compared to that estimated from static structural design loads alone.
2.3 Laminar-Flying-Wing Design Concepts
2.3.1 Greener by Design
In arriving at the performance estimates for a LFW, Green (2002) bases his evaluation on a number of parameters from the detailed conceptual BWB study of Bolsunovsky et al. (2001).
However, attempts are incorporated to allow for differences between the two configurations.
The additional weight and volume requirement of an integrated suction system is thought to favour a lower aspect ratio wing, and is consequently set at a value of 5. With a more
integrated flying wing, it is argued that the vortex induced drag factor should be lower, and closer to unity, therefore a value of 1.1 is assumed. The profile drag coefficient is taken as a quarter of that estimated for the the BWB, CD0 = 0.0026 — Green (2002) deems this value to be conservative based on the F-94A flight test results (see Sec. 2.1.1).
An overall propulsive efficiency typical of modern high bypass ratio turbofan engines is assumed at 37%. Green (2002) proposes that if the cruise Mach number is dropped from 0.85 to 0.80, and an unducted fan is utilised, an overall engine efficiency of 42% is achievable.
The LFW is taken to have a takeoff weight and span equal to that for the BWB. However, the wing loading is adjusted to account for a reduced aspect ratio and lower optimum cruise altitude.
The aircraft’s mission fuel is estimated using empirical correlations based on conventional aircraft, but modified to incorporate an increase of 7.5% takeoff weight to allow for suction hardware installation (based on an empirical correlation provided by Gregory (1961)), and a 10% reduction in structural weight associated with payload; however, potential weight savings with a lower aspect ratio are not accounted for.
At 2001 technology levels, Green (2002) estimates, for a LFW utilising turbofan technol-ogy, a reduction of around 50% on a baseline turbulent aircraft. Utilising UDF engines, this figure is projected to increase to around 60%. Allowing for 2050 technology improvements, this value further increases to around 70%.
2.3.2 Delta-Wing Aircraft with Suction
Denning et al. (1997) argue that an integrated delta-wing aircraft, with aspect ratio in the range 3 – 4, benefits from a low structural weight that can be traded off against a lowerL/D.
They add that it also permits a higher wing sweep, and hence a larger chord length for a given wing thickness (i.e. lower thickness-to-chord ratio), which then enables higher critical Mach numbers, better aeroelastic performance, and larger permissible internal volumes.
Denning et al. (1997) draw upon Lanchester’s minimum drag condition to show that:
CL∝√
CD0. This expression implies that a reduction in profile drag consequently leads to a reduction in cruise altitude and higher critical Mach numbers via a reducedCL; nevertheless, the aerodynamic efficiency will increase as L/D ∝ p
1/CD0. In contrast, Denning et al.
(1997) comment that the current aircraft design philosophy is to decrease induced drag by increasing aspect ratio as this permits higher values ofCL, requiring an increase in optimum cruise altitude and higher speeds for a given Mach number.
For a fixed aircraft geometry, and assuming: a constant cruise Mach number, optimum L/D, cruise-climb-datum cruise altitude, at fixed wing loading, of 40,000 ft, Denning et al.
2.3. LAMINAR-FLYING-WING DESIGN CONCEPTS
(1997) investigate the effect of an arbitrary profile drag reduction on altitude and non-dimensional engine thrust FN/δ (where δ is the ambient static pressure relative to that at sea level). They found that a 50% reduction in profile drag gives a 41% increase inL/D, an increase in ambient pressure by 41% (i.e. a reduction in cruise altitude to 33,000 ft), and a 50% reduction in non-dimensional engine thrust requirement. Denning et al. (1997) also document the impact that additional suction penalties, in terms of the suction coefficient, have on range performance. In Fig. 2.20 we see that the overall range performance is rather insensitive to the suction assumptions.
Figure 2.20: Laminarised aircraft — relative range performance — fixed integrated delta planforms. Assumptions: optimum L/D cruise (no allowances), cruise Mach number of 0.8, datum CD0 = 0.010 (no suction treatment), and CQ = 0.0005. (Taken from Denning et al.
(1997)).
2.3.3 The Handley Page HP117 Proposal
The basic design philosophy of the HP117 proposal, documented by Lee (1961), is focussed on the minimisation of direct operating costs. This parameter acts as a performance index for both transport efficiency and first cost. Economy of operation is achieved through the combination of LFC with the all-wing arrangement and a high cruise speed. However, some loss in aerodynamic/engine efficiency is accepted to reduce design/development expenses.
Four constraints are imposed on the design: 1) low cruise drag, 2) large internal volume, 3) high critical Mach number, and 4) good low-speed behaviour. The mission specification consists of: transatlantic flight at a cruise Mach number of 0.8, with a payload of 300 pas-sengers and 4500 kg of freight. (Lee (1961) comments that the true benefits of the LFW can only be realised with over 200 passengers and a range greater than 2000 nm.) The proposed configuration is shown in Fig. 2.21, and the design specification is summarised in Tab. 2.2.
Figure 2.21: Handley Page HP117 LFW configuration (from Green (2006)).
The design features a high degree of sweepback from the perspective of yielding a rea-sonable value for thickness-to-chord ratio to permit satisfactory passenger accommodation, whilst, combined with a low cruiseCL, allowing for a high cruise Mach number; however, Lee (1961) highlights that this is limited to 50◦ for fear of tip stall issues, and problems with the undercarriage design and adverse lateral stability characteristics.
2.3. LAMINAR-FLYING-WING DESIGN CONCEPTS
For a given payload, and hence cabin geometry, the wing area is minimised to reduce structural weight. It is argued that smaller wing areas also have the added benefit of higher wing loadings, lower cruise altitudes and, consequently, smaller engines. However, the upper value for wing loading is limited by low speed performance, unit Reynolds number, and the desire for a high critical Mach number.
Taper ratio is selected on the basis of passenger accommodation volume. However, the need for wingtip fins, which provide the necessary directional control, prevents the adoption of a triangular planform, whilst structural efficiency and passenger comfort considerations limit the feasibility of an untapered planform. The final taper ratio leads to a cabin volume 10 – 15% lower volume than the maximum possible.
Cabin volume reduces with aspect ratio; however, direct operating costs, assessed on the basis of structural and powerplant considerations, also reduce with aspect ratio. Therefore, a compromised aspect ratio is selected.
Table 2.2: HP117 aircraft design specification (from Lee (1961)).
Maximum thickness (m) 2.9
Root thickness-to-chord ratio 0.13 Outboard thickness-to-chord ratio 0.20 Tip thickness-to-chord ratio 0.15
Span (m) 40
Unit Reynolds number (m−1) 7.5×106
Sweep (degrees) 50
Mean chord (m) 14.5
Planform area (m2) 581
Aspect ratio 2.7
Taper ratio 0.33
Cruise lift coefficient 0.17
Mach number 0.80
Altitude (ft) 30,000
Velocity (m/s) 243
Allowable weight (kg) 140×103
Wing loading (N/m2) 2340
The cabin is integrated with the outer wing structure: the upper and lower panels are supported by the front and rear spars, and ribs. The suction surface is integrated within the structure. The structural elements are sized based on load conditions, and include: the skins;
the spars, including curved cabin pressure walls; and the floors. The rest of the aircraft is sized using standard weight estimation techniques available at the time for a conventional metal
alloy. In contrast to conventional aircraft, the report suggests that ground loads would be a more critical design condition than in flight, due to the distribution of structural and payload weight across the wing span. A structural weight fraction of 27% MTOW is estimated;
however, Lee (1961) comments that the cabin weight estimate may be optimistic.
A low cruise drag is achieved using slit suction: the resulting low profile drag permits a large wing area without a parasite drag penalty, and hence low span loading and induced drag.
Laminarised tip fins increase the effective aspect ratio, without any costs in terms of parasite drag. Consequently, a vortex induced drag factor of 0.943 is assumed. To avoid turbulence contamination downstream of areas not available for laminarisation, such as the cockpit windshield and doors, ‘clean-up’ slots are utilised, which enable a new laminar boundary layer to begin downstream.
The powerplant installation consists of six (very) low bypass-ratio Rolls-Royce RB.163 engines. One group of engines receive sucked boundary layer flow, whilst another group receive air from the clean-up slots, supplemented with contributions from ram intake. This scheme was selected as it had an advantage in terms of first cost benefit over a separate suction pump system. As prefigured by Lachmann (1955), Lee (1961) identifies takeoff thrust as the most significant engine design feature.
Lee (1961) gives the overall, effective, cruise L/D at 33. With a cruise lift coefficient of 0.17, a vortex-induced drag factor of 0.943, and an aspect ratio of 2.7, the induced drag coefficient CD,i is 0.0032, and therefore the profile drag coefficient (including pump drag) CD0 is 0.0020. Consequently, the low value for aspect ratio has compromised the aircraft’s aerodynamic efficiency. On the basis of the Handley Page cruise and climb performance studies, which assume suction availability for altitudes above 15,000 ft, the estimated fuel burn of the HP117 is around 18 g/pax.km.