Example Airframe Aerodynamics
Having established what may be achieved with CFD codes for sections and simple wings, we next apply these to a whole airframe, in this case for the Decode-1 UAV outer mold line already created using AirConics. We show the use of panel codes for aerodynamics and stability and RANS approaches for aerodynamics.
Figure 13.19 Experimental and computational lift and drag data for the Sivells and Spooner  wing and y+ for the k — a SST Harpoon mesh. Source: NASA.
Analyzing Decode-1 with XFLR5: Aerodynamics
When using XFLR5, only the lifting surfaces can be analyzed for lift and drag (though deflected control surfaces can be included by adding small trailing-edge distortions to the shapes of the airfoil sections where the controls are assumed to be). While the fuselage
Figure 13.20 XFLR5 model of the Sivells and Spooner wing.
can be included, the guidance supplied with the code advises against it; the tail booms and undercarriage details must be omitted completely. They can, however, be allowed in the drag calculations by being added as extra drag coefficients with associated areas. These can be taken from the previous spreadsheet analysis or from texts such as Hoerner . Figure 13.23 shows the resulting model built directly from the AirCONICS system and imported to XFLR5 via the medium of an XML file, while Figure 13.24 shows a set of polar plots for three flight speeds, for a mass of 15 kg and a fixed center of gravity (CoG) position 42.8 mm forward of the main wing quarter chord position (as taken from the previous spreadsheet analysis and with the CoG input to XFLR5 as 39.1 mm from the main wing leading edge). The plot of the imported geometry also shows the lift distribution in the cruise condition. Such data can be used to inform subsequent structural analysis such as sizing of the main spars.
Note that the previous spreadsheet-based approach to aerodynamic design was not based on any given airfoils: rather, typical airfoil lift and drag performance were assumed. By using XFLR5 and XFoil, it is possible to begin to understand the impact of actual section choices on the aircraft. Here, a NACA 23 012 section has been used for the main wing and 0212 for the elevator and fins. The main wing has zero setting angle and zero washout at the tips while an elevator setting angle of -2.85° has been chosen. Figure 13.24 shows that, in this configuration, at 30m/s, the aircraft has to fly at an AoA of 2.53° to achieve the required cruise Cl of 0.28 where the lift to drag ratio is comfortably greater than the value of 16 assumed in the spreadsheet. The chosen elevator setting angle allows the aircraft to fly stably at this AoA as can be seen by the fact that Cm is then zero. Moreover, the Cm value is positive at the zero lift AoA and the Cm curve has a negative slope, indicating that the design is stable in pitch, as expected.
This is not, however, an efficient configuration, as the fuselage would not be aligned with the direction of flight in the cruise condition, leading to extra drag. It is general practice to adjust the main wing and tail setting angles to produce a condition where both the required Cl and a zero Cm are achieved with the aircraft horizontal. To do this, it is useful to conduct a series of fixed-speed, fixed-AoA, and stability runs in XFLR5 with varying wing and elevator settings to achieve the required conditions. The stability calculation is particularly useful in this regard,
Figure 13.21 Experimental and computational lift and drag data for the Sivells and Spooner  wing with enhanced k — a SST Harpoon mesh of 76 million cells and y+ for the enhanced mesh. Source: NASA.
as it automatically searches for the speed and AoA where Cm is zero and reports these in the operating point. To begin this process, we first set the main wings to a setting angle of 2.53°
Figure 13.22 Pathlines and static pressure around the Sivells and Spooner  wing with enhanced k — a SST Harpoon mesh at 11° angle of attack. Source: NASA.
Figure 13.23 XFLR5 model of Decode-1 airframe as generated by AirCONICS with main wing setting angle of 0° and elevator setting angle of -2.85°, at an angle of attack of 2.6° and 30m/s. Note the use of cambered sections for the main wing and symmetrical profiles for the elevator and fins. The green bars indicate the section lift, with the tail producing downforce to ensure pitch stability.
(i.e., the AoA where a Cl of 0.28 was previously achieved) and reduce the tail setting angle by a similar amount. XFLR5 then suggests that the stable flight point would be at 30.08 m/s and at an AoA of -0.06°. A further iteration to the tail setting angle of -0.34° yields a flight point at 29.98 m/s and an AoA of -0.04°, which is the desired condition to three significant figures. Figure 13.25 shows the lift, drag, and moment polar plots of this final configuration.
Figure 13.24 XFLR5-generated polar plot for Decode-1 airframe as generated by AirCONICS with main wing setting angle of 0° and elevator setting angle of -2.85°, showing speed variations from 15to30 m/s. The black circles indicate flight at an angle of attack of 2.53° at which Cm is zero.
Figure 13.25 XFLR5-generated polar plot for Decode-1 airframe as generated by AirCONICS with main wing setting angle of 2.53° and elevator setting angle of -0.34°, showing speed variations from 15 to 30m/s. Note that Cl is 0.28 and Cm is zero at an angle of attack of 0° as required in the cruise condition.
Note here that at the assumed landing speed of 15m/s, the required Cl is 1.11, which is achieved at an AoA of approximately 9.25° where the lift/drag ratio is 17; XFLR5 will not solve above 12.2° for this configuration and flight speed, indicating that the stall angle is around this value, thus giving a stall margin during landing of some 25%. This margin may seem quite large, but in fact XFLR5 tends to be optimistic with regard to stall angles and lift/drag ratios, so this is a sensible value to aim for.
In Chapter 11 on spreadsheet-based design, care was taken to ensure static stability was obtained, and estimates of the impact of the wake of the main wing on the tail were used to do this. As just shown, with XFLR5 it is possible to check this analysis by studying the pitching moment coefficient curve. Note that the curve can be adjusted by altering the position of the CoG (this alters the slope), by changing the setting angle for the elevator (this moves the curve up and down on the plot), by moving the longitudinal position of the elevator relative to the main wing (and thus changing its lever arm, altering both slope and position of the curve), or by altering its vertical position (and thus changing the impact of the main wing downwash on the elevator). Of course, one is not free to set these arbitrarily; clearly, the longitudinal CoG must be realistic given the design in hand and large setting angles for the elevator will increase drag, both by directly increasing the drag of the elevator and also by requiring greater lift from the main wing. If an acceptable position for the CoG in front of the neutral point is not immediately achieved, it may be necessary to reposition heavy items within the aircraft; in the extreme case it may be necessary to lengthen the fuselage of the aircraft forward to do this. Alternately, if the design uses simple carbon-fiber spars for the tail booms, it may be easier to adjust the longitudinal position of the tail. Hopefully a sensible tail volume coefficient will have been adopted at the outset so that significant changes are not required. Figure 13.26 shows a set of polar plots for different CoG positions, tail length, and elevator setting angle for Decode-1.
-  We find that flap deflections beyond around ±5° cause XFoil to fail and this unfortunately limits the utility of thiscapability.
-  Further details on setting up and interpreting all the information generated by the XFLR5 stability calculation aregiven in the next section.
-  These results are compared with Fluent calculations and experimental data in the section on Fluent analysis ofDecode-1.