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17th Annual CFD Symposium CFD Division, August 11-12, 2015, Bangalore
Influence of Canopy shape on the supersonic drag of a generic Fighter Aircraft
Abhiram C, Neeraj S, Abdul R R,Libin G,Pathanjali.R.J+,Muralidhar.M++,Dr.Subhendhu S
Aeronautical Development Agency, Bangalore, India
Abstract
Present day multi role fighter aircrafts have a design flight condition optimized for high supersonic flow
regimes. Wave drag is the form of pressure drag that dominate the total aircraft drag in supersonic Mach numbers.
Improving the area ruling the aircraft geometry shows great promise in improving the supersonic wave drag. The
present study investigates the supersonic drag improvement studies carried out in a generic delta winged fighter
aircraft with changes in its fuselage. The area distribution for this delta winged fighter aircraft is improved by
modifying the canopy geometry for better supersonic drag. Various aircraft geometries with different canopy shapes
were analyzed in RANS solver CFD++. This best canopy geometry also resulted in improved trim drag and
improved supersonic performance of the aircraft.
Keyword: Canopy, Area rule, CFD, Wave drag, Supersonic
Introduction
Modern multi role fighter aircrafts are expected to perform equally good in subsonic as well as supersonic regimes.
At supersonic speeds, wave drag is the major component of aircraft drag. In order to achieve superior supersonic
performance, the wave drag of the aircraft needs to be reviewed in accordance with its area distribution and local
flow gradients. Exhaustive literature [Ref. 1,2,3,4,5] studies indicate that the supersonic drag can be reduced by
proper area distribution in the aircraft. Literature studies also indicate that the Sears-Haack distribution gives a
minimum wave drag for the given maximum area and length. Preliminary area rule studies [Ref.6] are carried out to
identify the scope for wave drag reduction. Subsequently the canopy region of the fuselage was modified and
studied in CFD++ to improve the supersonic drag of the aircraft. Various canopy geometries were tried out and the
best canopy shape constrained by the limitations of ejection seat and safety requirements, improves the area rule
significantly giving less supersonic wave drag. All CFD studies carried out is using the RANS solver CFD++
predominantly in the supersonic regime. The flight performance computations for various canopy geometries were
estimated using an In-House point mass performance model YAPP [Ref.7]
Basic Area Rule studies
The delta winged fighter aircraft considered for the present study has a cross sectional area distribution as in Fig 1.
The area distribution is obtained by taking normal sections of the aircraft geometry perpendicular to its longitudinal
axis to give an area rule applicable for Mach M=1. Initial Euler studies were carried out on an equivalent body of
revolution obtained from this area distribution in Fig.1. An equivalent Sears-Haack geometry based on the length
and maximum cross section area of the aircraft is also studied. Double SH is obtained by an asymmetric SearsHaack distribution with the maximum area shifted to coincide with the aircraft maximum cross section area. SH +
OCC is obtained by merging only the front half of Sears-Haack with the rear area distribution of the aircraft. The
delta benefit in wave drag is shown in parentheses for each of the area distributions. It can be seen that by improving
the first half of the area rule of the aircraft by Sears-Haack distribution, a wave drag benefit of 43 counts is obtained.
Subsequently IterI geometry [Fig.2] which was created to remove the sharp gradient in the front half of the area rule
keeping all other regions the same, gave a significant drag benefit of 40 counts. The sharp gradient in the first half of
the area rule occurs ahead of the intake near the canopy region. Based on these preliminary studies, it was decided to
modify the aircraft geometry near the canopy to improve the area gradient in the front half of the area rule.
Canopy modifications
Based on the preliminary area rule studies, geometric modification of the canopy was taken up to improve area rule.
The CFD analysis of Operational Clean Configuration (OCC) of the fighter aircraft is solved with hybrid
unstructured ICEMCFD grid in CFD++. All computations of the aircraft on this Reynolds Averaged Navier Stokes
(RANS) solver CFD++ is with Spalart-Allmaras turbulence model. Preliminary canopy geometries l1 and l2 (shown
+
[email protected], Pathanjali R J, ADA, Bangalore
[email protected], Muralidhar Madhusudan, ADA, Bangalore
++
in Fig.3) shows significant wave drag reduction to the level of IterI by improving the sharp gradient in the area rule.
A detailed flow analysis of the base aircraft configuration (base
(base-OCC) was carried out to understand
under
the local flow
gradients so that the
he subsequent canopy modifications can be optimally redesigned.
Fig 1: Area distribution
ibution for fighter aircraft,,
Sears Haack and double Sears Haack
aack bodies
Fig 2: Area distribution for fighter
aircraft and the effect of front balcony
Fig 3: Area distribution
stribution for fighter
aircraft and some modified geometries
The drag built-up at M=1.2 along the longitudinal axis (X) of the base aircraft geometry is shown in Fig 4. The Area
distribution along the length of the aircraft is give in the primary axis and its corresponding cumulative drag buildup
is shown in the secondary axis. It can be seen that there is a sharp increase in drag buildup in the first
fir half of the area
distribution exactly coinciding with the sharp gradient in the area rule around the canopy region. The reason for the
sharp
harp increase in drag buildup was analysed using its equivalent body of revolution as shown in Fig.5. The Mach
contours above the canopy indicate a sharp acceleration in accordance with the sharp change in the canopy
geometry. The supersonic flow over-acc
accelerates
elerates sharply (indicated by the dark red region above the canopy which
eventually ends in a shock, a little downstream. Similar acceleration gradients are observed in its equivalent body of
revolution geometry exactly at the region of sharp change in th
the canopy region.
Fig 4:: Area distribution for fighter aircraft, and its
drag built-up
up along the length of the aircraft
Fig 5: Flow pattern in Fighter aircraft and its
equivalent Body of Revolution geometry
Various canopy geometries were created and studied in CFD++ to improve the sharp gradient in the front half of the
area rule. The profile shape of few selected canopies and the final shape is shown in Fig. 6. The profile shape
shap of the
final canopy was decided based on the lowest drag profile from the CFD++ studies subject to the geometric
constraints of ejection
jection seat and safety ssystem and other manufacturing requirements. The Cp distribution on the top
generator line (Y=0 Plane, Y axis is perpendicular to the plane of the paper) for the base canopy and the modified
shape is shown in Fig.7. The profile shape of the geometry is shown in the primary Y axis for the base geometry
(blue) and final geometry (red). The Cp distribution on this top generator line (for the base and the final shape) is
plotted on the secondary axis in Fig.7.. Also the Cp distribution on the top generator line of the equivalent body of
revolution geometry [Axi-Sym] for the base and final canopy is shown in Fig.7.. It can be seen that in the region
where the canopy profile takes a sharp turn, there is sharp acceleration indicated by a steep dip in Cp for the base
configuration. Improving the canopy curvature has resulted in a lesser dip in the Cp contour for the final modified
17th Annual CFD Symposium CFD Division, August 11
11--12, 2015, Bangalore
Influence of Canopy Geometry on supersonic drag of a Delta wing Fighter Aircraft
Abhiram C, Neeraj S, Abdul R R, Libin G,, Pathanjali.R.J, Muralidhar.M, Dr.Subhendhu.
Dr.Subhendhu S
shape. The greater the dip in Cp, the stronger is the shock downstream for the base configuration as compared to the
final modified canopy. Similar trends in Cp gradients are observed in the values extracted from Body of Revolution
studies. This reduction in wave drag in the final shape can be attributed to two reasons, firstly the reduction in peak
acceleration on the top surface of the canopy, secondly the reduction in shock gradients slightly downstream of the
peak acceleration zone. The Final shape gives a supersonic drag benefit of approximately 12 counts. The increase in
canopy surface area has a marginal increase in skin friction drag. However improvement in the local flow gradients
in the subsonic regime reduces pressure drag in the modified canopy configuration and compensates for the increase
in skin friction. Therefore the reduction in subsonic drag is marginal in the modified canopy configuration. The
structural changes required to cater for the aerodynamic loads of the new canopy accounts for approximately 11kg.
The cross section area of the base canopy and the modified canopy shown in Fig.7 indicates the additional space
available behind the pilot. Approximately 50kg of additional fuel was estimated to be accommodated by rearranging
the inboards of the aircraft in the additional space created by the increase in the canopy profile.
Fig 6: Few canopy shapes studied in CFD++ and the final shape of the canopy
Fig 7: Cp distribution on the top surface line of canopy (left) and the cross section views of modified canopy
The drag buildup and its corresponding area rule for the modified canopy configuration is shown in Fig.8. It can be
seen that the crossover of the cumulative drag due to modified canopy occurs near the canopy region and continues
to remain less than the base configuration. The Mach contours from Euler CFD++ studies shown in Fig. 8 indicates
the extent of diffusion in velocity gradients near the modified canopy region. The peak Mach number reached on the
top surface of the modified canopy is diffused across a larger region, with a comparatively weaker shock
downstream (green region).
RANS studies for high AOA was carried out in CFD++ to understand the influence of canopy geometry on wing
flow characteristics. The wing vortex contours are not influenced by the change in canopy profile for 0o AOSS
(Angle of Side slip) as seen from Fig.9. The X vorticity contours are plotted for base and modified canopy
geometries and merged into a single image for ease of comparison. The minor vortex structures formed behind the
modified canopy changes significantly as compared to the base configuration. The Wing vortex streamlines also
confirm that the wing flow is least affected by the change in canopy shape. The modified canopy streamlines
however show a marked change in flow characteristics near the canopy region as compared to the base version. The
cross flow component present in base canopy is visibly reduced in the modified canopy streamlines.
17th Annual CFD Symposium CFD Division, August 11-12, 2015, Bangalore
Influence of Canopy Geometry on supersonic drag of a Delta wing Fighter Aircraft
Abhiram C, Neeraj S, Abdul R R, Libin G, Pathanjali.R.J, Muralidhar.M, Dr.Subhendhu. S
Fig 8: Drag buildup with the modified canopy and Mach contours from Euler CFD++ studies
Fig 9: Vortex contours on the aircraft with the base and the modified canopy
The change in fuselage characteristics ahead of the CG (Center of Gravity) has a marked impact on the pitching and
directional stability [Fig.10] of the aircraft with modified canopy. The smooth flow on the modified canopy surface
has improved the basic CMo (Zero lift pitching moment) and this delta is consistently carried over for a large range
of flight alpha until cross over happens near the extremum of the alpha envelope. The change in CMo due to change
in fuselage flow is favorable in terms of less trim drag for the modified canopy configuration. The upward shift in
CMo results in lesser control surface deflections for the modified canopy configuration and hence less trim drag.
Fig 10: Pitching and directional stability characteristics of base canopy and modified canopy
17th Annual CFD Symposium CFD Division, August 11-12, 2015, Bangalore
Influence of Canopy Geometry on supersonic drag of a Delta wing Fighter Aircraft
Abhiram C, Neeraj S, Abdul R R, Libin G, Pathanjali.R.J, Muralidhar.M, Dr.Subhendhu. S
The increase in the projected area of the fuselage (side view) ahead of the CG has a detrimental effect on the
dynamic stability characteristics of the configuration with modified canopy. The directional stability derivative
(Cnβ) has deteriorated in the modified canopy configuration due to less restoring force (Cn) of the fin. This
degradation in the directional stability of the modified canopy is observed throughout the alpha envelope and results
in a smaller cross over alpha as compared to the base version. This drawback due to reduction in usable alpha in the
modified canopy configuration is proposed to be overcome by a closed loop directional stability feedback control.
The present supersonic wave drag reduction due to modified canopy has a huge impact on the performance of the
aircraft. The wave drag reduction in the modified fighter is approximately 6% in supersonic region as compared to
the base configuration which translates to a 20% improvement in transonic accelerations. Reduction in wave drag
has improved the maximum level speeds of the modified fighter by 2%.
Conclusion
The supersonic wave drag characteristics of a generic fighter aircraft was studied initially based on its cross
sectional area distribution. Preliminary Sears-Haack shapes and equivalent body of revolution geometries indicated
significant scope in wave drag reduction. The canopy region was identified for shape modification to reduce drag
and various geometries were studied in RANS solver CFD++ constrained by seat & safety system, structural and
manufacturing requirements to arrive at the best canopy shape. The final modified canopy gave 6% reduction in
supersonic wave drag which translated to 20% improvement in transonic accelerations and 2% improvement in
maximum level speeds. The pitching stability characteristics have improved in the modified fighter which results in
less trim drag. The directional stability characteristics of the modified configuration has deteriorated resulting in
lower cross over alpha which is proposed to be overcome by closed loop feedback control. This modified geometry
also benefits with additional space in the aircraft behind the pilot that can be used to appropriately accommodate
more internal fuel.
Acknowledgement
Preliminary area rule studies carried out by Mr. Abhishek Kaushal as a precursor to this present study is
gratefully acknowledged.
References
1
Richard Seebass, "Supersonic Aerodynamics: Lift and Drag", RTO AVT Course on Fluid Dynamics Research on Supersonic
Aircraft, 1998, pp. 2.1-2.6.
2
Middleton W. D., Lundry J. L., "Aerodynamic Design and Analysis system for supersonic aircraft", NASA CR-2520, 1975.
3
Middleton W. D., Lundry J. L., Coleman R. G., " Aerodynamic Design and Analysis system for supersonic aircraft", NASA
CR-2521, 1975
4
Nimeesha B., Kuntawala, Jason E. and David W. Zingg, "Preliminary Aerodynamic Shape Optimization of A BlendedWing-Body Aircraft Configuration", 49th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace
Exposition, 2011.
5
John M. Morgenstern, Alan E. Arslan, "Mach Weighted Area Ruling For Supersonic Vehicles", United States Patent
Application Publication, US 2005/0121555 A1, 2005.
6
Pathanjali R. J., Muralidhar Madhusudan, " Aerodynamic Improvement Studies in V4.8.2 of LCA AF Mk2", Aeronautical
Development Agency, ADA/ARD&P/TN/061, 2015.
7
YAPP Performance Model, Dr.Koruthu, Private communication
17th Annual CFD Symposium CFD Division, August 11-12, 2015, Bangalore
Influence of Canopy Geometry on supersonic drag of a Delta wing Fighter Aircraft
Abhiram C, Neeraj S, Abdul R R, Libin G, Pathanjali.R.J, Muralidhar.M, Dr.Subhendhu. S