Reference · 46 terms

Glossary

Short, precise definitions of the words and symbols used in this module, each linked to the lesson where it is taught. Filter the list as you type, jump to a letter, or look a symbol up in Symbols at a glance.

Showing all 46 terms

A

Aerodynamic center: The point about which the pitching moment does not change with angle of attack: the quarter chord of an airfoil, and the quarter point of the mean aerodynamic chord of a subsonic wing.; See: Lesson 2, Lesson 6Related: Mean aerodynamic chord \(\bar c\), Pitching moment coefficient \(c_{m,c/4}\)
Airfoil: The cross-section of a wing in a plane parallel to the flight direction: a mean camber line with a thickness distribution laid on either side of it.; See: Lesson 1Related: Mean camber line, Thickness ratio \(t/c\)
Anhedral: Negative dihedral: the wing droops from the root. Used on high, swept wings, which have too much dihedral effect already.; See: Lesson 7Related: Dihedral \(\Gamma\), Dihedral effect
Aspect ratio \(A\): \(A = b^2/S\): how slender the wing is. Higher \(A\) cuts induced drag (\(K = 1/(\pi Ae)\)) and steepens the lift curve, but makes the wing heavier.; See: Lesson 4Related: Reference wing, Lift-curve slope \(\CLa\)

B

Boundary layer: The thin layer next to the surface in which viscosity slows the flow to rest at the wall. It starts laminar (thin, little friction, easily separated) and turns turbulent at transition (thicker, more friction, but it resists separation).; See: Lesson 2Related: Reynolds number \(Re\), Separation

C

Camber: The greatest height of the mean camber line above the chord, as a fraction of the chord. It sets the zero-lift angle and the pitching moment.; See: Lesson 1Related: Mean camber line, Zero-lift angle \(\alpha_{L0}\)
Center of pressure: The point where the lift alone would act, \(x_{cp}/c = 0.25 - c_{m,c/4}/c_l\). On a cambered airfoil it moves aft as the lift falls, so the fixed aerodynamic center is used instead.; See: Lesson 2Related: Aerodynamic center, Pitching moment coefficient \(c_{m,c/4}\)
Chord \(c\): The straight line from the leading edge to the trailing edge of an airfoil, and its length. Airfoil coordinates are fractions of it.; See: Lesson 1Related: Airfoil, Mean geometric chord \(\bar c_g\)
Critical Mach number: The flight Mach number at which the local flow over the airfoil first reaches Mach 1. The drag starts to rise a little above it.; See: Lesson 3Related: Drag-divergence Mach number \(\MDD\), Supercritical airfoil

D

Design lift coefficient: The lift coefficient an airfoil is designed to work best at; chosen near the wing's cruise \(C_L = (W/S)/q\), which for a jet's best range is \(\sqrt{\CDz/K}/\sqrt3\).; See: Lesson 3Related: Ideal lift coefficient \(c_{l,i}\), Korn equation
Dihedral \(\Gamma\): The upward angle of the wing from the root, seen from the front. Chosen from guidelines by wing position and sweep: about 5 to 7 degrees for a low unswept wing, anhedral for a high swept one.; See: Lesson 7Related: Dihedral effect, Anhedral
Dihedral effect: The roll away from a sideslip that gives lateral stability. Geometric dihedral, aft sweep (about \(1^\circ\) per \(10^\circ\)) and a high wing all add to it.; See: Lesson 7Related: Dihedral \(\Gamma\), Wing position
Drag bucket: The narrow range of lift coefficients over which a laminar-flow airfoil keeps laminar flow and has much lower drag.; See: Lesson 2Related: Laminar-flow airfoil, Design lift coefficient
Drag-divergence Mach number \(\MDD\): The Mach number at which the transonic drag rise becomes steep (\(\dd\CD/\dd M = 0.1\)). A subsonic jet cruises just below it.; See: Lesson 3Related: Korn equation, Critical Mach number

E

Elliptic span loading: A lift per unit span that varies as \(\sqrt{1 - (2y/b)^2}\): the loading with the least induced drag for a given span. A taper ratio near 0.45 comes close.; See: Lesson 6Related: Taper ratio \(\lambda\), Schrenk's approximation

I

Ideal lift coefficient \(c_{l,i}\): The lift coefficient at which the flow meets the leading edge smoothly; \(\pi A_1\) in thin-airfoil theory. About 0.26 for the NACA 2412.; See: Lesson 2Related: Design lift coefficient, Thin-airfoil theory
Incidence \(i_w\): The angle between the root chord and the fuselage axis, set so the fuselage is level in cruise: \(i_w = \alpha_{L0} + C_L/\CLa\). About \(2^\circ\) for light aircraft, \(1^\circ\) for transports.; See: Lesson 7Related: Lift-curve slope \(\CLa\), Zero-lift angle \(\alpha_{L0}\)

K

Korn equation: \(\MDD = \kappa_A/\cos\Lambda - (t/c)/\cos^2\Lambda - c_l/(10\cos^3\Lambda)\), with \(\kappa_A = 0.87\) for conventional and 0.95 for supercritical airfoils.; See: Lesson 3Related: Drag-divergence Mach number \(\MDD\), Supercritical airfoil

L

Laminar-flow airfoil: An airfoil (such as the NACA six-series) shaped to keep the boundary layer laminar farther aft, giving a drag bucket near its design lift coefficient.; See: Lesson 1Related: Drag bucket, NACA four-digit airfoil
Leading-edge radius: The radius of the circle that fits the nose of an airfoil; \(1.1019\,t^2c\) for the NACA four-digit family. A sharp nose stalls abruptly.; See: Lesson 1Related: Airfoil, Stall
Lift-curve slope \(\CLa\): The rise of lift coefficient per radian of angle of attack: \(2\pi\) for a thin airfoil, less for a finite wing (Helmbold's formula; Prandtl's lifting line, \(2\pi A/(A + 2)\), agrees only for slender wings), more at high subsonic Mach number, less with sweep.; See: Lesson 4Related: Aspect ratio \(A\), Thin-airfoil theory

M

Mean aerodynamic chord \(\bar c\): \(\bar c = \tfrac23c_r(1 + \lambda + \lambda^2)/(1 + \lambda)\), at \(\bar y = (b/6)(1 + 2\lambda)/(1 + \lambda)\) from the centerline. The reference length for pitching moments; CG positions are quoted in % MAC.; See: Lesson 6Related: Aerodynamic center, Mean geometric chord \(\bar c_g\)
Mean camber line: The curve halfway between the upper and lower surfaces of an airfoil. Thin-airfoil theory computes lift and moment from it alone.; See: Lesson 1Related: Camber, Thin-airfoil theory
Mean geometric chord \(\bar c_g\): \(S/b\): the chord of a rectangular wing of the same span and area. Not the same as the mean aerodynamic chord unless the wing is rectangular.; See: Lesson 4Related: Mean aerodynamic chord \(\bar c\), Aspect ratio \(A\)

N

NACA four-digit airfoil: NACA \(mptt\): maximum camber \(m\)% of the chord at \(p\) tenths of the chord, thickness \(tt\)%, from two camber parabolas and a standard thickness distribution.; See: Lesson 1Related: Camber, Thickness ratio \(t/c\)
Normal Mach number \(M_n\): \(M\cos\Lambda\): the Mach number of the flow normal to the sweep, the only part the airfoil of a swept wing feels (simple sweep theory).; See: Lesson 5Related: Sweep \(\Lambda\), Drag-divergence Mach number \(\MDD\)

P

Pitch-up: The nose-up pitch of an aft-swept wing whose tips stall first: they are aft of the center of gravity, so losing their lift raises the nose further.; See: Lesson 5Related: Sweep \(\Lambda\), Twist and washout
Pitching moment coefficient \(c_{m,c/4}\): The moment about the quarter chord, per \(qc^2\) per unit span. Zero for a symmetric airfoil, negative (nose down) for positive camber; thin-airfoil theory gives \((\pi/4)(A_2 - A_1)\).; See: Lesson 2Related: Aerodynamic center, Thin-airfoil theory

R

Reference wing: The trapezoid that matches the real wing outboard, extended through the fuselage to the centerline. Its area \(S\) is the one used in every coefficient.; See: Lesson 4Related: Aspect ratio \(A\), Mean aerodynamic chord \(\bar c\)
Reynolds number \(Re\): \(\rho Vc/\mu\), the ratio of inertial to viscous forces. It sets the boundary layer, and so the drag and the stall: compare airfoils only at the design's Reynolds number.; See: Lesson 2Related: Stall, Drag bucket

S

Schrenk's approximation: The span loading of an unswept, untwisted wing taken as the average of its chord distribution and an elliptic one of the same area; \(c_l/C_L = (c + c_e)/(2c)\) shows where it stalls first.; See: Lesson 6Related: Elliptic span loading, Taper ratio \(\lambda\)
Separation: Where the boundary layer runs out of momentum against a rising pressure and leaves the surface; behind it the flow near the wall runs backward. Separation causes the stall, and much of the drag at high lift.; See: Lesson 2, Lesson 2Related: Boundary layer, Stall
Stall: The loss of lift when the flow separates at high angle of attack. Thick airfoils stall from the trailing edge (gently), moderate ones from the leading edge (abruptly), thin ones with a long bubble.; See: Lesson 2Related: Leading-edge radius, Thickness ratio \(t/c\)
Subsonic leading edge: A leading edge swept behind the Mach cone, \(\Lambda_{LE} > 90^\circ - \arcsin(1/M)\), or \(M\cos\Lambda_{LE} < 1\): the flow normal to it is subsonic, so it can be round.; See: Lesson 5Related: Sweep \(\Lambda\), Normal Mach number \(M_n\)
Supercritical airfoil: An airfoil with a flat top and aft camber that weakens and delays the shock, so it can be thicker at the same Mach number (\(\kappa_A = 0.95\) in the Korn equation).; See: Lesson 1, Lesson 3Related: Korn equation, Drag-divergence Mach number \(\MDD\)
Sweep \(\Lambda\): The angle of a chord line (leading edge, quarter chord, …) behind the spanwise direction. It lowers the normal Mach number, at a cost in lift, stall behavior and weight.; See: Lesson 5, Lesson 6Related: Normal Mach number \(M_n\), Pitch-up

T

T-tail: A horizontal tail on top of the fin, clear of the wing wake and rear-mounted engines; the fin is heavier, and a deep stall can blanket the tail.; See: Lesson 8Related: Tail volume coefficient, Tail arm \(L\)
Tail arm \(L\): The distance from the wing's quarter-MAC to the tail's: about 60% of the fuselage length with a nose engine, 50 to 55% with wing engines, 45 to 50% with rear engines.; See: Lesson 8Related: Tail volume coefficient, Mean aerodynamic chord \(\bar c\)
Tail volume coefficient: \(c_{HT} = L_{HT}S_{HT}/(\bar c_WS_W)\) and \(c_{VT} = L_{VT}S_{VT}/(b_WS_W)\): the tail's area times its arm, compared with the wing's. Statistics give a first tail size.; See: Lesson 8Related: Tail arm \(L\), T-tail
Taper ratio \(\lambda\): \(c_t/c_r\). About 0.45 brings an unswept wing close to elliptic loading; swept wings use 0.2 to 0.35. Too much taper overloads the tips.; See: Lesson 6Related: Elliptic span loading, Schrenk's approximation
Thickness ratio \(t/c\): The greatest thickness of an airfoil as a fraction of its chord. Thicker saves structure and holds fuel, but lowers the drag-divergence Mach number.; See: Lesson 1, Lesson 3Related: Korn equation, Stall
Thin-airfoil theory: An inviscid theory of the camber line: \(c_l = 2\pi(\alpha - \alpha_{L0})\), with \(\alpha_{L0}\), \(c_{m,c/4}\) and \(c_{l,i}\) from integrals of the camber slope.; See: Lesson 2Related: Zero-lift angle \(\alpha_{L0}\), Pitching moment coefficient \(c_{m,c/4}\)
Twist and washout: Twist sets the sections at different angles along the span; washout (tip lower than root) unloads the tips so the stall starts inboard. Typically 0 to 5 degrees; it suits one lift coefficient best, and costs a little induced drag on a strongly tapered wing.; See: Lesson 7Related: Pitch-up, Incidence \(i_w\)

W

Wing position: High wings suit cargo aircraft and ground clearance, low wings airliners (gear and spar under the cabin), mid wings fighters (least interference drag).; See: Lesson 7Related: Dihedral effect, Dihedral \(\Gamma\)
Winglet: A near-vertical surface at the wingtip that reduces induced drag much as more span would; used where the span is limited.; See: Lesson 7Related: Aspect ratio \(A\), Elliptic span loading

Z

Zero-lift angle \(\alpha_{L0}\): The angle of attack at which an airfoil gives no lift: negative for positive camber, about \(-1^\circ\) per 1% of camber at 40% of the chord. The airfoil's zero-lift line, through the trailing edge, lies \(|\alpha_{L0}|\) above the chord at the nose.; See: Lesson 2Related: Thin-airfoil theory, Camber

Symbols at a glance

Met a symbol in a lesson or in another textbook and not sure what it stands for? Find it here, then follow the link to its entry.

Symbols used in this module
Symbol	Meaning	Entry
\(c,\ t/c,\ m,\ p\)	Chord, thickness ratio, camber and its position	NACA four-digit airfoil
\(\alpha_{L0},\ c_{m,c/4},\ c_{l,i}\)	Zero-lift angle, moment about the quarter chord, ideal lift coefficient	Thin-airfoil theory
\(Re,\ \MDD,\ \kappa_A\)	Reynolds number, drag-divergence Mach number, Korn airfoil factor	Drag-divergence Mach number \(\MDD\)
\(S,\ b,\ A,\ \bar c_g\)	Reference area, span, aspect ratio, mean geometric chord	Aspect ratio \(A\)
\(\CLa,\ K,\ e\)	Lift-curve slope, induced-drag factor, Oswald factor	Lift-curve slope \(\CLa\)
\(\Lambda,\ M_n\)	Sweep, normal Mach number	Sweep \(\Lambda\)
\(\lambda,\ c_r,\ c_t\)	Taper ratio, root and tip chords	Taper ratio \(\lambda\)
\(\bar c,\ \bar y\)	Mean aerodynamic chord and its distance from the centerline	Mean aerodynamic chord \(\bar c\)
\(i_w,\ \Gamma\)	Incidence, dihedral	Incidence \(i_w\)
\(c_{HT},\ c_{VT},\ L\)	Tail volume coefficients, tail arm	Tail volume coefficient

Statistical values are from D. P. Raymer, Aircraft Design: A Conceptual Approach, Chapters 4 and 6; check them against your edition.