A good aerodynamics design is vital to the success of a supersonic aircraft, with
the two main objectives as the ability of supersonic flight and the reduction of take off
gross weigh. The design of the wing must provide sufficient lift to meet performance
requirements, while the overall aircraft design is required to minimize drag to reduce total
The Sentinel is designed for the highest performance with the lowest possible cost
in terms of total fly away cost or the amount of fuel burned.
5.2. Wing Design
5.2.1. Airfoil Selection
Although the geometry of a thin airfoil does not make a significant difference
during supersonic flight, it does greatly influence the subsonic performance of the
aircraft. The NACA 64-206 airfoil was chosen for the Sentinel based on the airfoil
thickness and the availability of airfoil wing tunnel data. The airfoil has a 6% chord
thickness to reduce the negative shockwaves effect for supersonic performance while the
slight camber enables the airfoil to produce lift even at zero angle of attack. The NACA
64 series was reverse designed for the minimization of drag comparing to the other
NACA airfoils. The geometry layout of the airfoil can be seen in Fig. 5.1.
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Figure 5.1. Non-dimensional geometry of NACA 64-206 airfoil.
5.2.2. Wing Geometry and Performance
The Sentinel was designed with a swept wing with an aspect ratio of 3.25, a taper
ratio of 1/3, and leading edge sweep angle of 50 degree. Fig. 5.2. illustrates the wing
geometry and Mach cone generated by supersonic flight.
Lengthwise dimension [ft]
-40 -20 0 20 40
Spanwise dimension [ft]
Figure 5.2. Graphic illustration of the wing geometry with a Mach cone at Mach 2.2.
The wing was strategically positioned to be completely engulfed by the Mach cone,
which results in reduced flow velocity over the wing. This setup is advantageous since a
wing in a subsonic flow produces approximately twice the lift as a wing in supersonic
flow according to the conical flow theory .
Effects such as downwash, wing twist, and taper ratio all have influence on the
transformation between airfoil aerodynamic data and the aerodynamic properties of a
three dimensional finite wing. DATCOM charts provided by Raymer were used to
predict the performance of the wing under various flow conditions. This data was
compiled into Fig. 5.3, providing a correlation between maximum lift coefficients at
various Mach numbers, although it should be noted that the maximum lift produced at
higher Mach numbers depend more on the structural strength of the wing.
0 0.5 1 M 1.5 2 2.5
Figure 5.3. Maximum lift coefficient versus Mach number.
5.3. Drag Prediction
Drag prediction is essential to the success of designing a supersonic aircraft. A
reduction in drag produced by the aircraft would result in an increased access thrust, thus
greater climb rate and higher overall performance; it would also reduce the fuel
consumption and decreased the take off gross weight required for the mission.
Drag forces on an aircraft are formed by three main components: the lift induced
drag which is caused by the pressure difference between the front and back of the wing,
the parasitic drag which is mainly consist of skin friction and turbulence formed behind
the aircraft due to the shape or “form” of the components, and the wave drag which is
caused by the formation of shock wave during transonic and supersonic flight. Fig. 5.4
compares the wave drag of the Sentinel with various historical aircrafts.
0 0.5 1 1.5 2 2.5
Figure 5.4. Wave drag comparison between the Sentinel and F-106, F-4 and B-70.
A combination of parasitic drag and wave drag at various Mach number and
altitude is shown in Fig. 5.5. Calculations based on the Boeing estimation approach yield
a drag divergent Mach number of 0.894; at which point the wave drag becomes a
significant part of the overall drag produced by the aircraft. Data from this plot and the
lift induced drag were combined to estimate the thrust required for steady level flight.
0.018 40000 20000
0.5 1 1.5 2 2.5
Figure 5.5. Parasitic drag and wave drag at various altitudes [ft] and Mach numbers.
Another essential parameter in determining the aircraft performance is the lift to
drag ratio, usually obtained from the drag polar plot of an aircraft. This ratio is essential
in predicting the optimal altitude, fuel consumption, best cruise velocity and bestloiter
velocity. It is also used to predict the maximum climb rate velocity. Fig. 5.6 displays the
drag polar diagram of the Sentinel.
0 0.1 0.2 0.3 0.4
Figure 5.6. Drag polar of the Sentinel at 40,000 ft.
5.4. Comparison with Delta Wing
Delta wing configurations are common among modern supersonic aircrafts such
as the Eurofighter, Dassault Rafael, and Saab Gripen. The delta wing is known to have
good transonic and supersonic performance due to the highly swept leading edge, which
also contributes to the formation of leading edge vortex (LEX). The resulting LEX
effectively provides additional lift production at higher angles of attack. In order to
justify the decision to employ a swept back wing instead of delta wing, drag analysis and
CL,max data from DATCOM for both wing types were compiled and compared in Fig.
5.7 a) and b).
Swept Wing 35000ft
Delta Wing 35000ft
0 0.5 1 1.5 2 2.5
0 0.5 1 1.5 2 2.5
Figure 5.7. The comparison between swept wing of Sentinel and a delta wing of the
same wing area for a) drag forces at various Mach number and b) Maximum lift
coefficient at various Mach number using DATCOM data
. As illustrated in the above plots, the delta wing produced slightly less drag at
transonic and supersonic regime comparing to the swept back wing. However, the
maximum lift coefficient of the delta wing is significantly lower than that of a swept back
wing. A swept back wing was chosen based on the above statement in addition to the
known property of having a higher aspect ratio at a given wing loading value, thus
producing less induced drag.
5.5. Lift Distribution
In order to study the effect of downwash, as well as to decide on the setup of
structural components to withstand the lift generated by the wing, an approximation of
lift distribution along the wing is required. There are various methods that can be used to
solve for the lift distribution of a tapered swept wing, with the most influential one as the
modified lifting line theory proposed by Weissinger . However, due to numerical
difficulty, it was decided to use a simpler approximation suggested by Schrenk which
uses equation 5.1.
dL 1 dCl y
q c( y ) c 1 ( 5.1),
dy 2 d 4 b
where c(y) is the variation of chord length with respect to spanwise location. The
resulting half span lift distribution of the wing is shown in Fig. 5.8.
CL / y
0 0.2 0.4 0.6 0.8 1
Normalized y location
Figure 5.8. The spanwise lift distribution of the wing at zero degree angle of attack.
5.6. High Lift Devices
Extra lift is needed during take off and landing as the aircraft is traveling at a
slower velocity, which would require a larger CL value to stay aloft. High lift devices
were introduced in order to generate adequate lift for take off, landing, and high g turns.
The Sentinel employed 70% span leading edge slats and Fowler type flaps that cover the
inner 30% of the exposed wing area, similar to the setup illustrated in Fig. 5.9. A Fowler
type flap was chosen since it generates the highest lift among the flap types and can be
reasonably equipped on an airfoil of 6% chord thickness. The increase in lift from the
usage of high lift devices under various conditions is demonstrated in Table 5.1.
Figure 5.9. F-18 with full span leading edge slat.
Table 5.1. Contribution of high lift devices for various maneuvers.
Altitude SL 35000 35000
Mach # 0.2 0.7 1.2
Maneuver Landing Turn1 Turn2
c'/c flap 1.1 1.001 1.025
c'/c slat 1.025 1.00025 1.0125
ΔCDo 0.0523 0.0476 0.0487
ΔCDi 0.0163 0.0142 0.0147
ΔCL,max,total 0.5522 0.5157 0.5258
5.7. Interdisciplinary Trade Study
The instantaneous turn rate is one of the most demanding requirements set by the
RFP. The aircraft stall limit and 7 g structural limit was plotted to obtain the corner speed
at which the maximum instantaneous turn rate occurs. A trade study of various flap
configurations was done to optimize the flap setting as discussed in section 3.2.
A good aerodynamics design is essential for a successful supersonic aircraft. The
analysis presented above has illustrated that the Sentinel is optimized for supersonic
operation as well as subsonic maneuvers. Various obstacles were encountered and
overcame during the design process, demonstrating the TSN design team motto “Think,
 Bertin, J. “Aerodynamic for Engineeers”, 4th edition, Princeton Hall Publication,
 Raymer, D. “Aircraft Design: A Conceptual Approach”. 3rd edition, AIAA, 1999.
 Weissinger “Lift Distribution of Swept-Back Wing” NACA TM 1120, 1947.
 Schrenk “A Simple Approximation Method for Obtaining the Spanwise Lift
Distribution” NACA TM 948, 1940.
 Yves Fauconnier “J-5005 about to land after a nice in flight display at Nancy air
show. June 2002.” [http://perso.wanadoo.fr/aeromil-
yf/F18l%205005%20landing.jpg. Accessed 4/12/06.]