| Term | Definition |
|---|
| standard atmosphere | 1. Temperature distribution |
| lapse rate | alpha, the (change in temperature/change in altitude): alpha=dT/dh=-6.51 deg. Celcius/km |
| stratosphere lapse rate | lapse rate=0 |
| troposphere | h < 11 km, alpha= dT/dh |
| stratosphere | 11 < h < 25 km |
| R (for air) | 1716.5 ft^2/sec^2*Rankine OR 287.84 m^2/sec^2*Kelvin |
| reynolds number | pho*V*l/mu OR V*l/v (v=kinematic viscosity) |
| kinematic viscosity | dynamic viscosity/density |
| Reynold's Number | (normal pressures * inertia forces)/(viscous forces) |
| velocity and pressure | interdependent forces |
| bernouli's equation | can't be applied to flows where the effects of viscosity are appreciable such as the boundary layer |
| total pressure | dynamic pressure + static pressure |
| pitot static tube | connected to an airspeed indicator, static pressure on outside dynamic pressure on inside |
| airspeed indicator | always calibrated for standard sea level conditions |
| relative density | sigma, density/density sea level |
| viscosity shearing state | due to fluid momentum transport between adjacent layers having different layers |
| laminar flow | momentum transport due only to molecular diffusion |
| shear stress | resistance to motion along the x-axis |
| turbulent flow | reynolds number becomes greater than 12,000 to 14,000 |
| momentum transport | due primarily to bulk momentum transfer in the turbulent eddies |
| boundary layer edge | 0.99*Ambient Velocity |
| velocity profiles | tend to remain similar |
| velocity gradient at the wall | decreases as x increases and the skin friction stress also decreases as x increases |
| local-skin friction coefficient | force / (reference stress * reference area) |
| skin friction versus distance | on a flat plate, change in velocity/change in height decrease as does skin friction with x |
| integrated skin friction coefficient | introduces the average skin friction coefficient, cf |
| laminar boundary layer | 1. Parabolic velocity profiles |
| turbulent boundary layer | a. Velocity profiles are fuller. b. boundary layer thickness is greater, (height turbulent > height laminar) c. Velocity gradient is greater, thus Skin Friction Stress is greater e. A laminar sub layer always exists near the surface for turbulent boundary layers. f. |
| Skin Friction Drag Coefficient (One side only) | cd = Drag Force/ (dynamic velocity*reference area) |
| mixed flow | both laminar and turbulent flow on the same surface |
| natural or spontaneous transition | reynolds number range: 2*10^5 <= Retr <= 2 * 10^6 |
| no slip condition | velocity at the wall is equal to velocity of the wall |
| adverse pressure gradient | dp/dx > 0. in this the flow momentum decreases i.e. the flow velocity decreases as the static pressure increases |
| static pressure | approximately constant across the boundary layer, dp/dy = 0 |
| fluid layers near the wall | less momentum than outer layers |
| outer fluid layers | more momentum than fluid layers near the wall |
| adverse pressure gradient | greatest relative effect on the inner layers and brings fluid to rest before the outer layer |
| fuller turbulent boundary layer | goes further in an adverse pressure gradient |
| fuller turbulent boundary layer | delays boundary layer separation |
| fuller turbulent boundary layer | cost is increased skin-friction drag (golf ball) |
| displacement thickness | effectively uncambers the airfoil |
| displacement thickness | reduces lift and causes a wake thickness which increases drag |
| inviscid fluid | normal pressures all balance |
| inviscid fluid | no form drag as a result of normal pressures balance happens in this |
| inviscid flow pressure distribution | constant total pressure throughout |
| inviscid flow pressure distribution | stagnation points at leading edge and trailing edge (non-cusped airfoil) |
| inviscid flow pressure distribution | stagnation pressures the same at both stagnation points |
| inviscid flow pressure distribution | no net drag acting on the airfoil |
| viscous flow pressure distribution | total pressure loss in viscous boundary layer |
| viscous flow pressure distribution | stagnation points near the leading edge and the sharp trailing edge |
| viscous flow pressure distribution | pressures at the trailing edge stagnation point is approximately equal to the free-stream static |
| viscous flow pressure distribution | region of adverse pressure gradient over aft portion of airfoil on upper surface tends to separate the boundary layer |
| viscous flow pressure distribution | drag is now present |
| drag source | skin friction drag |
| drag source | the total pressure loss in the boundary layer causes a |
| boundary layer separation | produces a drag increase and a lift decrease (stall) |
| changing the effective airfoil shape | interfere with the pressure and the upper surface trailing edge |
| changing the effective airfoil shape will | interfere with the generation of large negative pressures on the upper surface near the trailing edge |
| boundary layer separation | can produce an increased pressure drag and decreased lift |
| turbulent boundary layer | tends to delay separation and delay stall |
| turbulent boundary layer | a cost is skin friction drag |
| turbulent boundary layer | desirable at low speed conditions |
| vortex flow | M << 1.0 |
| vortex flow | characterized by (two dimensional) concentric circular streamlines |
| irrotational flow | v= constant/R |
| real vortex | has a non-potential rotational core |
| potential vortex | everywhere zero except at the origin R=0 |
| circulation | line integral about some closed path of the product of the v-component along the path and the distance along the path |
| circulation | sum of vorticity along a closed path |
| biot savart law | vortex filament in 3-d |
| doubly infinite vortex | vortex filament extends from - infinity to +infinity |
| kutta joukowski theorem | force per unit span acting on a right cylinder of any cross-section equal to rho*V*gamma and acts perpendicular to Vinfinity, no drag |
| total flow | vector sum of the component velocities |
| lift | perpendicular to ambient velocity |
| sectional lift | lift force per unit span or width of the lifting body |
| bound vortex | fixed to a given set of spacial coordinates or to a given body |
| free vortex | remains fixed to a given set of fluid particles, moves with the particle |
| kutta condition | requires smooth flow off a sharp trailing edge, finite angle, stagnation point, cusp, on upper and lower surface |
| kutta condition | removes the (impossible) velocity discontinuity at the sharp trailing edge |
| circulation | this term is formed by the viscosity acting in the boundary layer |
| hemholz vortex theorem | vortex strength remains constant along the length of a vortex filament |
| hemholtz vortex theorem | filament cannot end in the fluid |
| hemholtz vortex theorem | without rotational and external forces an initially irrotational fluid remains irrotational |
| starting vortex | the fluid around a body at rest has no circulation |
| starting vortex | body put into motion the fluid will have some circulation |
| turbulent boundary layer | a. Velocity profiles are fuller |
| turbulent boundary layer | b. boundary layer thickness is greater, (height turbulent > height laminar) |
| turbulent boundary layer | c. Velocity gradient is greater, thus Skin Friction Stress is greater |
| laminar boundary layer | 2. Aerodynamically smooth |
| turbulent boundary layer | e. A laminar sub layer always exists near the surface for this |
| standard atmosphere | 2. Standard Sea-level properties assumed |
| standard atmosphere | 3. Equation of state is taken as p=rho*R*T |
| standard atmosphere | 4. Atmospheric hydrostatic equation dp=-rho*g*dh |
| laminar boundary layer | 3. no streamwise pressure gradient |
| chord line | straight line joining ends of camber line |
| camber line | line midway between upper and lower surface |
| thickness | profile height perpendicular to chord line |
| leading edge radius | radius of circle tangent to upper and lower surfaces with center on a line tangent to camber line at leading edge |
| (V infinity) relative wind | free stream velocity in 2-D case |
| alpha | geometric angle of attack |
| L prime | sectional lift force (lift force per unit span) always perpendicular to relative wind |
| D prime | sectional drag force, always parallel to relative wind |
| N prime | sectional normal force, always perpendicular to chord line |
| C prime | sectional chord force, always parallel to chord line |
| M prime | sectional pitching moment is denoted as: |
| reynolds number increase | lift curve moves along same path, but has a higher angle of attack |
| reynolds number increase | drag curve moves down and to the right, parallel to the old drag curve |
| center of pressure | location (usually on chordwise portion) of moment center about which the moment is zero |
| aerodynamic center | location (xac) of moment center about which the pitching moment is independent of angle of attack |
| profile drag | subsonic 2-D drag |
| profile drag | has two parts, (pressure or form drag) + (skin friction drag) |
| pressure or form drag | due to viscous pressure losses and boundary layer separation |
| skin friction drag | due to viscous shearing at the body's surface |
| cd profile | cd pressure or form + cd friction |
| flat plat normal to flow | cd profile = cd pressure |
| flat plate parallel to flow | cd profile = cd friction |
| thickness problem solution | distributing sources and sinks along the chordline and solving for the source sink distribution which yields the given shape |
| thickness on airfoil | contributes nothing to the lift or moment on the airfoil |
| 1. laminar separation bubble | laminar flow |
| 2. laminar separation bubble | stagnation point |
| 3. laminar separation bubble | transition point |
| 4. laminar separation bubble | separation bubble |
| 5. laminar separation bubble | turbulent reattachment point |
| 6. laminar separation bubble | turbulent flow |
| 7. laminar separation bubble | downstream velocity |
| 1. model irrotational until | shocks |
| model irrotational flow | once viscous effects occure one can no longer do this |
| 3. model irrotational until | heat loss or gain |
| low angle of attack | sectional lift = sectional normal force |
| low angle of attack | sectional drag = sectional chord force |
| C prime | does not include skin friction, only includes pressure force |
| separation | can happen laminar and is not synonymous with transition |
| delays separation | higher reynolds number |
| viscous forces | this force determines separation |
| p infinity | static pressure |
| transition | reynolds number influences this |
| transition | surface characteristics or roughness influences this |
| transition | airfoil shape influences this |
| transition | angle of attack influences this |
| transition | freestream turbulence influences this |
| transition | noise influences this |
| transition | suction blowing influences this |
| 2.0V | In a potential flow with a free-stream velocity of V, the velocity on the top of a right circular cylinder is |
| the decrease in the normal velocity gradient at the wall due to boundary layer thickening | the decrease in the local skin friction as the flow proceeds from the leading edge of a flat plat at zero angle of attack is due to |
| 1.0 | in incompressible flow, the maximum positive pressure coefficient that can be achieved, for example, a forward stagnation point |
| decrease | increasing the freestream reynolds number generally causes the drag coefficient of an airfoil to |
| boundary layer transition | can be "natural' due to the growth of instabilities in the flow, or by means of a laminar separation bubble |
| zero for 2-D potential flows | the drag predicted by inviscid flow theory is |
| decrease pressure drag | dimples on a golf ball |
| subsonic profile drag | composed of pressure drag and viscous drag |
| characterized by reverse flow | two-dimensional boundary layer-separation is: |
| negative | for an airfoil with positive camber, the zero-lift angle of attack is |
| a function of angle of attack | the center of pressure is |
| transonic drag rise becomes significant | the divergence mach number is the free-stream mach number for which |
| the lift curve slope is always predicted to be 2*pi | from thin airfoil theory |
| the local flow first reaches sonic conditions at some point on the body | the critical mach number corresponds to the flow condition where |
| subsonic profile drag | composed of pressure drag and viscous drag |
| thin airfoil theory | allow the boundary conditions of the problem to be linearized |
| aerodynamic center | the chordwise point at which the airfoil pitching moment coefficient is independent of lift coefficient |
| kutta condition | a body with a sharp trailing edge will have smooth flow off that trailing edge |
| 0.127/degree | mach number is 0.2, the lift curve slope is 0.112, what is lift curve slope when it is m=0.5 |
| laminar drag bucket | shifted to higher conditions with increased camber |
| unpowered high lift device | approach will cause area to increase |
| unpowered high lift device | approach will cause circulation to increase |
| unpowered high lift device | approach will cause boundary layer control |
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