Validation of Cp on the L51F07 Configuration

ZEUS subsonic and supersonic steady pressure results (invicsid and viscous) for the L51F07 Wing-Body Configuration are compared to wind tunnel results.

Comparison of Wing Cp, M=0.9, AoA=2°

Comparison of Body Cp, M=0.9, AoA=2°

Comparison of Wing Cp, M=1.2, AoA=2°

Comparison of Body Cp, M=1.2, AoA=2°

Lift, Moment and Drag on L56A18 Wing-Body-Tail Configuration

Comparison of Normal Force Coefficient M = 0.80, 0.90, 0.94, 1.03

Comparison of Coefficient of Moment M = 0.80,0.90,0.94,1.03

Comparison of Coefficient of Drag M = 0.80,0.90,0.94,1.03

Bent-Nose Compact Kinetic Energy Missile (CKEM) with Wrap-around Fins

• Nine blocks of overset meshes to model the missile body and eight wrap-around fins
• Block 1: Missile Body
• Blocks 2-9: Eight wrap-around fins
• CPU time for each angle of attack case is approximately 10 minutes with one CPU

Mesh on Y-Z Plane

3D View of Mesh

Lift and Moment at Mach 0.6

Steady and Unsteady Cp on LANN Wing

M=0.822, α=0.6° Aspect Ratio: 7.92 Taper Ratio: 0.4 Swept Angle: 25° Twist from root to tip: 2.6° ~ -2.0° Supercritical Airfoil with max t/c = 12% Pitch mode at reduced frequency = 0.102

Steady Cp at Four Span Stations

Unsteady Cp (Real Part)

Unsteady Cp (Imaginary Part)

S⁴T Wind Tunnel Flutter Model

5 Blocks Block 1: Aircraft, 133x78x69 Block 2: Trail, 55x26x50 Block 3: Canard, 75x21x61 Block 4: Inbord Engine, 73x24x36 Block 5: Outboard Engine, 69x23x41

Angle-of-Attack

Trailing Edge Flap

Horizontal Tail

• Interactive process for solving the trim variables such as α, β, p, q, and r as well as aircraft accelerations.
• Capable of dealing with dertermined trim system as well as over-determined trim system (more unknown trim variables than trim equations).
• Generation of flight loadsand output of NASTRAN FORCE and MOMENT bulk data cards for subsequent detailed stress analysis.

Joined-Wing at M=0.8 and Load Factor = 2.0g

F-16 with Under-Wing Stores at M=0.95 and Load Factors = 5.0g

• Capable of generating frequency-domain generalized aerodynamic forces and computing flutter boundaries using the g-method.
• Transient response analysis for LCO predictions.

Flutter Analysis of Goland Wing

LCO Analysis of Goland Wing

24 Mesh Blocks Used to Model the Whole Aircraft with Under-wing Stores

Splined Structural Modes on the Aerodynamic Mesh

F-16 Tip Launcher Acceleration at AoA = 2.2° and Re=4.6x10⁶

• Flutter model was tested in NASA Langley Transonic Dynamic Tunnel (TDT) with heavy gas.
• Between M=0.77 to 0.82, two flutter modes were found; the Wing/Nacelle mode (17Hz) and the wing tip (22Hz).
• The Wing/Nacelle mode has low unstable damping (hump mode).
• >Dynamic pressure can be continuosly increased without damaging the model until encountering the wing tip flutter mode.
• Beyond M=0.82, the unstable damping of the Wing/Nacelle becomes high which can lead to the destruction of the wing due to flutter. For this reason, the wing-tip flutter mode condition cannot be reached.

• 12 natural modes are included in the flutter analysis.
• The infinite plate spline method (for the wing and pylon) and thin plate spline method (for the fuselage and Nacelle) are used to transfer the modes from the structural grid to the aerodynamic grid.

Two Blocks of Overset Mesh are used:

• Block 1: Fuselage, Wing, and pylon, 122x52x61
• Block 2: Engine Nacelle, 61x41x20

• The 17Hz hump mode predicted by ZEUS correlates with the low unstable damping of the Wing/Nacelle mode observed in the TDT.
• The flutter boundary of wing-tip mode predicted by ZEUS agrees well with TDT data while ZAERO linear aerodynamics largely over-predict the flutter boundary.