Since you asked for an explanation appropriate to an non-specialized audience, maybe this will do: "A Physical Description of Flight; Revisited" by David Anderson & Scott Eberhardt. It is a revision of the earlier "A Physical Description of Flight" (HTML version).

From Stick and Rudder by Wolfgang Langewiesche, page 9, published 1944:

The main fact of all heavier-than-air flight is this: the wing keeps the airplane up by pushing the air down.

It shoves the air down with its bottom surface, and it pulls the air down with its top surface; the latter action is the more important. But the really important thing to understand is that the wing, in whatever fashion, makes the air go down. In exerting a downward force upon the air, the wing receives an upward counterforce--by the same principle, known as Newton's law of action and reaction, which makes a gun recoil as it shoves the bullet out forward; and which makes the nozzle of a fire hose press backward heavily against the fireman as it shoots out a stream of water forward. Air is heavy; sea-level air weights about 2 pounds per cubic yard; thus, as your wings give a downward push to a cubic yard after cubic yard of that heavy stuff, they get upward reactions that are equally hefty.

That's what keeps an airplane up. Newton's law says that, if the wing pushes the air down, the air must push the wing up. It also puts the same thing the other way 'round: if the wing is to hold the airplane up in the fluid, ever-yielding air, it can do so only by pushing the air down. All the fancy physics of Bernoulli's Theorem, all the highbrow math of the circulation theory, all the diagrams showing the airflow on a wing--all that is only an elaboration and more detailed description of just how Newton's law fulfills itself--for instance, the rather interesting but (for the pilot) really quite useless observation that the wing does most of its downwashing work by suction, with its top surface. ...

Thus, if you will forget some of this excessive erudition, a wing becomes much easier to understand; it is in the last analysis nothing but an air deflector. It is an inclined plane, cleverly curved, to be sure, and elaborately streamlined, but still essentially an inclined plane. That's, after all, why that whole fascinating contraption of ours is called an air-plane.

This answer is nothing more than a variation of Sklivv's answer. I simply wish to discuss some quantitative ideas following from Sklivv's answer and discuss what I understand (from an aerospace engineering friend) to be a common conceptual mistake - that the application of "mere surface effects" and "application of Bernoulli's principle" is wrong. These "mere surface effects and Bernoulli's principle" follow from Sklivv's idea as I hope to make clear. Everything in aeroplane physics begins and ends with "aeroplanes thrust air downwards, so the air thrusts aeroplanes up". This answer is written to be understandable to someone like I who know nothing about fluid dynamics - aside from:

1. The mathematically elegant and thoroughly enjoyable 2D problems tackled with complex variable theory (see Finding Stagnation Points from the complex potential );

2. That I know that there is Clay Mathematics prize up for grabs for anyone who can prove existence of, or give counterexample against, the existence of smooth, globally well defined solutions to the Navier-Stokes equations;

3. That aerospace engineering colleagues and friends tell me experimental proof is still queen in this field: most real fluid dynamics involving aeroplane flight heavily leans on phenomenological models tuned by experiment.

I shall answer by taking these points up in turn.

Experiment is Queen

From a particular experimental point of view, there is no mystery why aeroplanes fly. Rather, the better question, in my opinion, is "how do they control the inevitable huge lift forces on them to make the latter stably lift in a constant, vertical direction?"

This experimental view is as follows: think of the Beaufort Scale and other scales used by meteorologists to get across the practical meaning of their wind and other warnings: for example the Fujita Scale for tornadoes and Tropical Cyclone Category Systems, which describe in practical terms the effects of storms of various intensities.

Now I understand flying regulations forbid commercial jetliners from flying at slower than $300mathrm{km,h^{-1}}$ before their final approach to the runway. Think about $300mathrm{km,h^{-1}}$ airspeed in terms of the scales I have just spoken of: this is an F4 tornado, category 5 cyclone and is well off the 12-class Beaufort scale. Buildings and structures of any shape the size and weight of fully laden aeroplanes are torn up and borne into the sky or utterly torn down and destroyed. There is NO shortage of lift from a $300mathrm{km,h^{-1}}$ relative airspeed to hold almost anything the size and weight of fully laden commercial jetliner up: at these airspeeds, almost anything of this size and weight and lighter flies. At least it does so fleetingly: if it isn't designed like an aeroplane, as it moves its attitude changes and so does the direction of ram pressure: it is then likely to be flipped over and dashed catastrophically onto the ground. Put simply: almost everything flies at this airspeed, but only very special things do so stably.

Simple Mathematical Models

We can do a back of the envelope estimation of ram pressure in this case: see my drawing below of a simple aerofoil with significant angle of attack being held stationary in a wind tunnel. I'm going to put some numbers to Sklivvz's description:

Lets suppose the airflow is deflected through some angle $theta$ radians to model an aeroplane's attitude (not altitude!) on its last approach to landing or as it takes off, flying at $300mathrm{km,h^{-1}}$ airspeed or roughly $80mathrm{m,s^{-1}}$. I have drawn it with a steep angle of attack. Air near sea-level atmospheric pressure has a density of about $1.25mathrm{kg,m^{-3}}$ (molar volume of $0.0224mathrm{m^{-3}})$. The change in momentum diagram is shown, whence the change in vertical and horizontal momentum components are (assuming the speed of flow stays roughly constant):

$$Delta p_v = p_b sintheta;quadquadDelta p_h = p_b ,(1-costheta)$$

At the same time, the deflecting wing presents an effective blocking area to the fluid of $alpha,A,sintheta$ where $A$ is the wing's actual area and $alpha$ a scale factor to account for the fact that in the steady state not only fluid right next to the wing is distrubed so that the wing's effective area will be bigger than its actual area. Therefore, the mass of air deflected each second is $rho,alpha,A,v,sintheta$ and the lift $L$ and drag $D$ (which force the engines must afford on takeoff) must be:

$$L = rho,alpha,A,v^2,(sintheta)^2;quadquad D = rho,alpha,A,v^2,(1-costheta), sintheta$$

If we plug in an angle of attack of 30 degrees, assume $alpha = 1$ and use $A = 1000mathrm{m^3}$ (roughly the figure for an Airbus A380 wing area), we get a lifting force $L$ for $rho = 1.25mathrm{kg,m^{-3}}$ and $v = 80mathrm{m,s^{-1}}$ of 200 tonne weight. This is rather less than the takeoff weight of a fully laden A380 Airbus (which is 592 tonnes, according to the A380 Wikipedia page) but it is an astonishingly high weight just the same and within the right order of magnitude. As I said, experiment is Queen here. We see that the wing's effective vertical cross section is bigger than the actual wing by a factor of 2 to 3. This is not surprising at steady state, well below speed of sound flow: the fluid bunches up and the disturbance is much bigger than just around the wing's neighbourhood. So, plugging in an $alpha = 3$ (given the experimental fact that the A380 can lift off at 592 tonnes gross laden weight), we get a drag $D$ of 54 tonne weight (538kN) - about half of the Airbus's full thrust of 1.2MN, so this ties in well with the Airbus's actual specifications, given there must be a comfortable margin to lift the aeroplane out of difficulty when needed.

In these F4 / C5 grade winds (and up to three times faster in normal flight), we see therefore there simply isn't any shortage of lift. The aeronautical engineering problem is more about keeping this plentiful lift stably directed upwards and allowing the aeroplane to hold a steady attitude and keep any torques arising from lift nonuniformity from flipping the plane over.

As the aeroplane picks up speed, the ram pressure calculated above is proportional to the square of the airspeed (see my answer to Drag force at high speeds ), so that at full speed the effect more than accounts for the drop in air density and the shallower angle of attack - we cannot make this downwards ram pressure without overcoming the much greater horizontal hindwards component - drag - so that is important to fly with low angle of attack for good fuel efficiency.

Refining the Mathematical Model

It is important to heed that the above description in terms of momentum difference between incoming air and the downwash begotten by the wing is exactly the same physics as the "more popular" descriptions given in terms of the Bernoulli equation and the integration of pressure around the wing. This is easy to see: the Navier-Stokes equation(See the Wikipedia page for the derivation of the Navier-Stokes equation ), is a very simple application of nothing more than Newton's second and third laws to infinitessimal volumes of fluid, notwithstanding the lack of knowledge about its fundamental mathematical properties (as bespoken by the Clay Mathematics Millenium Prize's unclaimed status: I love the Navier-Stokes equation- such a simple, readily grasped idea so baldly just an embodiment of Newton's laws, yet throwing up profound mysteries that show us scientists how little we yet know about the World). The steady state Navier Stokes equation for a perfect, incompressible fluid is (here $vec{v}$ is the steady state velocity field and $p$ the scalar pressure field):

$$(vec{v}cdot nabla) vec{v} = nabla left(frac{|vec{v}|^2}{2}right) + nablawedge(nablawedgevec{v}) = -nabla p$$

which gives $nablaleft(p + frac{|vec{v}|^2}{2}right) = 0$ or $p + frac{|vec{v}|^2}{2} = text{const}$ for an irrotational flow ($nablawedgevec{v} = vec{0}$) when integrated along the integral curve of $vec{v}$, i.e. a streamline. Or, alternatively, we can argue in a more first principles way in this simple case: the force on a infinitessimal volume is $-nabla p$ and the acceleration of a particle on the streamline is, by application of the Serret-Frenet formulas (here $s$ is the arc length along the streamline through the particle and $kappa$ the path's curvature):

$$mathrm{d}_t (v hat{mathbf{t}}) = mathrm{d}_s v times mathrm{d}_t s, hat{mathbf{t}} + v,mathrm{d}_s(hat{mathbf{t}}),mathrm{d}_t s=v,mathrm{d}_s v, hat{mathbf{t}} - kappa,v^2,hat{mathbf{n}}=mathrm{d}_s left(frac{v^2}{2}right), hat{mathbf{t}} - kappa,v^2,hat{mathbf{n}}$$

whence, on applying $vec{F} = m vec{a} Rightarrow -nabla p ,mathrm{d}x,mathrm{d}y,mathrm{d}z = rho,vec{a},mathrm{d}x,mathrm{d}y,mathrm{d}z$, we get:

$$-nabla p = rho left(mathrm{d}_s left(frac{v^2}{2}right), hat{mathbf{t}} - kappa,v^2,hat{mathbf{n}}right)$$

which again yields $p + frac{|vec{v}|^2}{2} = const$ when integrated along a streamline (here we can see the sideways (normal to streamline) centripetal force $-v^2,hat{mathbf{n}} / R$ given by the wonted $v^2/R$ formula). So we can (and will, below), for example, apply the Theorem of Blasius to calculate lift, and be assured it is no more than a quantification of Sklivv's idea that "aeroplanes thrust air downwards, so the air thrusts aeroplanes up". The pressure difference between the upper and lower surface of a wing exists because the wing is pushing the air down, not a separate phenomenon. Often one hears that the Bernoulli principle applied to wings is wrong: this is not true. There is a fallacy (to be discussed below) as shown by experiment (and, hand-wavingly, by theory) in the wonted demonstration of lift using Bernoulli's principle, but the idea is basically sound, as it must be from its derivation from the Navier-Stokes equation and Newton's laws shown above.

A Joukowsky Aerofoil Calculation and Errors in Wonted Application of Bernoulli's Principle to Wings

We look at a 2D calculation of lift by Bernoulli's principle, or, equivalently, by application of the Theorem of Blasius. The common misconception here is that airflows split at the wing's leading edge and two neighbouring particles will reach the wing's lagging edge at the same time, so that the upper particles must fare the curved surface at higher speeds and therefore the pressure on the upper wing surface is less. Actually, the upper path particles are sped up much more than this explanation implies and reach the wing's lagging edge well before their lower-path-faring neighbours. See this wonderful video from the University of Cambridge, particularly at about 50 seconds in. This fact shows that the circulation $oint_Gamma vec{v}cdotmathrm{d}vec{r}$ around the wing's surface $Gamma$ is nonzero, a fact which we intuitively expect from simple theory (as shown below) and which is amply confirmed in experiment: see the video, or go to the end of a runway of a large airport on a damp day so that you can let big commercial jetliners fly over you at about 50m height (take your earmuffs). On a damp day, you will see vortices breaking off the wings' outer edges, will see them swirling in the damp air for many seconds in the aeroplane's wake and, if you take you hearing protection off after the aeroplane has pass, you will hear the vortices crackling in the air, sounding a little like waves washing on the beach. This is much more fun than it sounds when your children are badgering you to do such a thing and, from the sights and sounds, I learnt heaps more from doing it that I thought I would. Even though the following calculation has an air of theoretical soundness and "first principles" to it, it is important to understand that it too is an experimental model: the circulation is forced into our description, motivated by the confirmation of the former's existence by experiment. The Kutta-Joukowski Condition (see Wikipedia page for Kutta Condition) as well as the Wikipedia page for the Kutta-Joukowski Theorem is little more than an ad-hoc experimentally motivated fix: it is simply this. When we model the flow with a Joukowski aerofoil (described below), there is a sharp, lagging edge on the wing. This begets a singularity with unphysical, infinite velocties. However, by postulating and choosing the right circulation in the flow, we can put a stagnation point at the lagging edge, thus cancelling the singularity, regularising our solution and also forcing the experimentally observed condition that there is only ever one stagnation point at the wing's leading edge, never elsewhere.

Another way to look at this experimentally motivated condition is well explained in this answer to the Physics SE question Does a wing in a potential flow have lift?. An irrotational, inviscid, incompressible flow cannot alone lift a wing. We add circulation to "fudge" a compensation for this theoretical lack: viscosity is "nature's way of enforcing the Kutta-Joukowsski condition".

So we begin with the complex variable method (see the Wikipedia page for "Potential flow" in the section "Analysis for two-dimensional flow" to study a potential flow i.e. irrotational ($nabla wedge = vec{0}$) velocity field $vec{v}$ with a potential $psi$ such that $vec{v} = -nabla psi$ that is also incompressible (continuity equation $nablacdot vec{v} = nabla^2 psi = 0$). See also the Physics SE questions Finding Stagnation Points from the complex potential ).

The main method here is to use the Joukowski transform:

$$omega(z,,s_z,,s_omega) = frac{ s_omega }{2}left(frac{z}{ s_z } + frac{ s_z }{z}right)$$

to map the potential flow corresponding to a spinning, offset cylinder (see the NASA page “Lift of a Rotating Cylinder" ) into the flow around the image of this cylinder under the Joukowsky transform. The truly weird Flettner Aeroplane actually used spinning cylinders rather than wings to fly successfully. The Joukowsky transform maps the circle $|z| = s_z$ onto the real axis between the points $omega = pm s_omega$ in the $omega$-plane; this section of the real axis between $omega = pm s_omega $ is then the branch cut for the inverse Joukowski transform. The Joukowsky transform is a two to one mapping, and the branches of the inverse Joukowski transform map the whole $omega$-Riemann sphere (if we define the stereographic projection so that $|z| = s_omega $ is the $omega$-Riemann sphere’s equator) separately to the inside and outside of the circle $|z| = s_z$ in the $z$-plane (which outside and inside can be thought of the Northern and Southern hemispheres of the $z$-Riemann sphere, if the stereographic projection is chosen so that the circle $|z| = s_z$ is the $z$-Riemann sphere’s equator). The $omega$-Riemann surface is made by slitting two copies of the Riemann sphere along the branch cut and stitching the edges together, to get a double cover of genus nought for the $omega$-Riemann sphere. For this problem, I define the branch cut as slightly differently from the real axis section between the $pm s_omega$, I define it as the path:

$$operatorname{Im}(omega) = h cosleft(frac{pi}{2} operatorname{Re}(omega)right)$$

between the two branch points with an adjustable height parameter $h$, for reasons that will become clear.

The radius $r$ of the spinning cylinder radius is chosen so that cylinder surface passes through the point $z=+s_z$, which is the image of one of the branch points in the $omega$ plane. This achieves the sharp edge that becomes the lagging edge of our aerofoil.

The complex potential for the spinning cylinder is:

$$Omega(z) = v ,e^{-ialpha},left(z- deltaright) + frac{r^2 ,v, e^{+ialpha }}{z- delta } + i,a,logleft(z - delta right)$$

where $alpha$ is the angle of attack, $delta = delta_r + i,delta_i$ is the offset and $r$ is the radius of the cylinder steeped in a uniform flow which converges to $v$ metres per second along the positive real axis, as $ztoinfty$. The logarithm and dipole terms put a branch point and pole at the cylinder’s centre, so the flow is perfectly valid outside and on the cylinder. $a$ is the circulation. If we let $phi$ stand for the angular co-ordinate labelling the cylinder’s edge, there are two stagnation points on the cylinder with angular co-ordinates $phi_pm$ where $mathrm{d}_z Omega(z) = 0$, i.e. when:

$$e^{i,(phi_pm - alpha)} = -ifrac{a}{2,v,r}pmsqrt{1-left(frac{a}{2,v,r }right)^2} = expleft(-arcsinfrac{a}{2,v,r }right)$$

Now, we map this flow to the $omega$ plane and apply the Theorem of Blasius to the image of the offset circle so as to work out the lift on this image. The image can be plotted with the Mathematica command:

$$small{mathrm{P[delta_r_, delta_i_] := ParametricPlot[{Re[omega[delta_r + i delta_i + sqrt{(1 - delta_r)^2 + delta_i^2} Exp[i theta]], Im[omega[delta_r + i delta_i + sqrt{(1 - delta_r)^2 + delta_i^2} Exp[i theta]]}, {theta, 0, 2 pi}]}}$$

and the result is drawn below in the $omega$-plane for $s_z = s_omega = 1$, $delta_r = -0.1$, $delta_i = 0.3$ (i.e. the spinning circle offset so that its centre is at $-0.1+i,0.2$ and with a radius $r = sqrt{(1 - delta_r)^2 + delta_i^2}$ so that its image passes through the branch point $omega = +s_omega = 1$ in the $omega$-plane:

Now we come to the crucial Kutta-Joukowski postulate, an experimental "fudge". The sharp edge on the aerofoil above would normally map the flow in the $z$-plane so that there were an unphysical infinite velocity at this sharp point. In practice, it is seen in wind tunnel tests that the streamlines stay tangent to the upper surface, and that there is one stagnation point at the wing's leading edge (intuitively the air "crashes" here) and no other stagnation points on either the top of bottom of the wing. Sometimes there is a small region of turbulence around the wing's lagging edge (as in the University of Cambridge video) (i.e. the incompressible potential flow model fails here) or the flow peels smoothly off the lagging edge. The way we achieve effects similar to experiment and "renormalise" our solution is to add the right amount of circulation $a$ to the flow so that one of the stagnation points on the spinning cylinder is mapped to the sharp edge (the branch point at $omega = +s_omega$) in the $omega$-plane: the stagnation thus cancels the otherwise unphysical infinite velocities there and "regularises" our solution. With the radius of the cylinder chosen as $r = sqrt{(1 - delta_r)^2 + delta_i^2}$, it can readily be shown from the equation above for the stagnation point positions that the circulation needed is:

$$a = 2 v,delta_i cosalpha + 2,v,(1-delta_r) sinalpha$$

This then is the wholly experimentally motivated Kutta-Joukowski condition. It is motivated by the knowledge that circulation is observed aroung wings, there is experimentally only one stagnation point on the leading edge of the wing and the fact that the right amount of circulation can reproduce these experimentally seen results.

When this is done, the Blasius theorem lift calculation done around the transformed Joukowski aerofoil in the $omega$-plane is:

$$begin{array}{lcl}D_ell - i,L_ell &=& frac{i,rho}{2}oint_{Gamma_omega} (mathrm{d}_omega Omega)^2 ,mathrm{d} omega &=& frac{i,rho}{2}oint_{Gamma_z} (mathrm{d}_z Omega)^2 frac{1}{mathrm{d}_z omega},mathrm{d} z &=& -pi,rho Sigma[,mathrm{residues,of,},(mathrm{d}_z Omega)^2 frac{1}{mathrm{d}_z omega},mathrm{at,poles,within,}Gamma] &=& -4,pi,i,rho,a,v,e^{-i,alpha}end{array}$$

where $Gamma_omega$ is the Joukowski aerofoil and $Gamma_z$ the transformed aerofoil (i.e. the spinning cylinder). So there is no lift without circulation. It's worth stating again:

An irrotational, inviscid, incompressible flow cannot alone lift a wing. We add circulation to "fudge" a compensation for this theoretical lack: viscosity is "nature's way of enforcing the Kutta-Joukowsski condition".

Now we substitute the Kutta-Joukowski condition to get:

$$D_ell + i,L_ell = 8,pi,i,rho,v^2,left(delta_i,cosalpha + (1-delta_r),sinalpharight) frac{s_z^2}{s_omega} e^{+ialpha}$$

We now need to scale the velocities so that the relative airspeeds are equal in the $omega$- and $z$-planes.

The above is the force per unit length (in direction normal to the page) on the wing and its direction is the direction in the $omega$-plane. We have:

$$limlimits_{omegatoinfty} left(mathrm{d}_omega Omega(omega(z))right) = limlimits_{ztoinfty} left(mathrm{d}_zOmega(omega(z))right) limlimits_{omegatoinfty} left(mathrm{d}_omega zright) = 2 ,e^{-ialpha} v frac{s_z}{s_omega}$$

so we need $s_omega = 2$ and $s_z = 1$, then $delta$ will be a dimensionless parameter defining the offset of the $z$-plane cylinder as a fraction of its radius. But now the $omega$-plane planform width of the wing is 4 units. Moreover, the above calculation yields the force per unit length (normal to the 2D flow). So we divide the result for $s_omega = 2$ and $s_z = 1$ by 4 and then scale up by the total wing area to get the total force on the wing. Furthermore, we need to we rotate the flow in the sketch below so that the incoming flow is horizontal (i.e. in the direction of the aeroplane's relative air velocity) in the $omega$-total force on the wing above becomes:

$$D + i,L = pi,i,rho,v^2,A,left(delta_i,cosalpha + (1-delta_r),sinalpharight)$$

We witness the d'Alembert paradox: the perfect flow cannot model the drag. Now let's put some numbers in. If we put $delta = 0$, then the wing is simply the straight branch cut between $omega = pm 1$, so we have a version of the calculation I began with but now refined to take account of the full flow pattern. With $alpha = 0.3$ (a little less than 20 degrees), $rho = 1.25mathrm{kg,m^{-3}}$, $v=80mathrm{m,s^{-1}}$ and $A = 850mathrm{m^2}$, we get $L=643mathrm{tonne}$, pretty near to the Airbus's fully laden takeoff weight. If we chose the parameters $delta_i = 0.2$, $delta_r =-0.1$ to give a wing shape that does not seem too fanciful for a jetliner wing with the lagging edge flaps wound fully out for takeoff and landing (see the plot below) we get about 1200 tonnes lift for our $300mathrm{km,h^{-1}}$ airspeed. Clearly this is optimistic and the overreckonning arises from the assumption of equal effectiveness of whole wingspan, whereas the tips will clearly not be well modelled by 2D flow. Not all of the wings will work as modelled, thus the $A$ in this formula is somewhat less than the planform area. What the flow model does show (see below), however, is that the effective vertical cross section presented to the incoming air is much greater than the tilted area $A ,sintheta$ assumed in the very simple model at the beginning of my answer. At steady state, a considerable cross section of air both above and below the vertical cross section is bent downwards and contributes to the effect "aeroplanes thrust air downwards, so the air thrusts aeroplanes up" described in Sklivv's answer.

Now, to plot the complete transformed flow in the $omega$-plane, we must use the inverse Joukowski transform. To do this successfully, one must use the right branches of the inverse transform in the right co-ordinate patches. For Mathematica, which puts the branch cut for the square root function along the negative real axis (the namespace std::sqrt in Microsoft Visual C++ puts it along the positive real axis), we define the following chart functions, which are particular branches of the inverse transform:

$$zeta_1(omega) = frac{s_z}{s_omega}left(omega- i sqrt{omega-s_omega},sqrt{-left(omega+s_omegaright)}right)$$ $$zeta_2(omega) = frac{s_z}{s_omega}left(omega+ i sqrt{omega-s_omega},sqrt{-left(omega+s_omegaright)}right)$$ $$zeta_3(omega) = frac{s_z}{s_omega}left(omega- sqrt{omega^2-s_omega^2}right)$$ $$zeta_4(omega) = frac{s_z}{s_omega}left(omega+ sqrt{omega^2-s_omega^2}right)$$

and then the following Mathematica commands will plot the full flow:

$$small{mathrm{Omega[z_,,delta_,,v_,,r_,,a_,,alpha_,,s_]:= v,e^{-i,alpha}left(frac{z}{s}-deltaright) + frac{r^2,v,e^{i,alpha}}{frac{z}{s}-delta} + i,a,Logleft[frac{z}{s}-deltaright]}}$$ $$small{mathrm{G[z_,,delta_r_,,delta_i_,,alpha_]:=Omegaleft[z,,delta_r+i,delta_i,,1,,sqrt{(1-delta_r)^2 + delta_i^2},2,delta_i Cos[alpha] + 2,(1-delta_r),Sin[alpha],,alpha,,1right]}}$$

$$small{mathrm{S[delta_r_, delta_i_, alpha_, h_, c_] := Show[ContourPlot[ Im[If[(Abs[x] < 1 ) wedge (y > 0) wedge (y < h, Cos[pi x/2]), G[zeta_1[x + i y], delta_r, delta_i, alpha]], If[x < 0, G[zeta_3[x + i y], delta_r, delta_i, alpha]], G[zeta_4[x + i y], delta_r, delta_i, alpha]]]]], {x, -2, 2}, {y, -2, 2}, Contours to c, MaxRecursionto 2, PlotPoints to 300, AspectRatio to 1], P[delta_r, delta_i, {Black, Thick}]]}}$$

where $mathrm{P}[]$ is the parametric plot command above used to plot the aerofoil. The above use of the branch functions works for $delta_r < 0$: other branches are needed for correct results when $delta_r > 0$. The parameter $h$ bends the branch cut so that it bows upwards and stays inside the aerofoil, thus allowing the branches of the inverse Joukowsky transform to plot the mapped cylinder flow properly. Drawn below is the outcome from the command $mathrm{S[-0.1, 0.2, 0.2, 0.2, 100]}$, i.e. the flow around the wing for an angle of attack of 0.2 radians, the circle offset parameters of $-0.1 + 0.2,i$, a bow in the branch cut so that $h=0.2$. Witness the branch cut inside the aerofoil below and also how far from the wing's surface its effect stretches. The effective vertical component of the wing's area that is presented to the flow is clearly much greater than the actual vertical component of the wing's area, so the factor of 2 to 3 scaling in the A380 Airbus lift as reckoned by the simple fluid deflexion calculation seems highly plausible and unsurprising.

Lastly, to come the full circle, here is an animation to be found on web pages "Irrotational plane flows of an inviscid fluid" at the University of Genoa's environmental engineering department; see http://www.diam.unige.it/~irro/. The animation shows the progress of fluid particles for the Joukowski aerofoil flow, illustrates the assertion that the flow above the wing traverses the wing much more quickly than the flow underneath and lastly, shows very well the main thesis that "aeroplanes thrust air downwards".

Without going into the excellent and detailed mechanics explaining reaction lift that others have provided for this answer, I just want to say that contrary to popular belief/high school physics textbooks, airplanes do not fly solely on account of Bernoulli's principle. According to Walter Lewin's excellent "For the Love of Physics":

"Bernoulli's principle accounts for 20% of an airplane's lift, the rest is provided by reaction lift."

Walter Lewin also poses an insightful question if planes really fly due to the equal transit theory and Bernoulli's principle (they do not!).

"...then how do planes fly upside down?"

Consider the velocity field of the particles in the air mass in a 2D projection of the X(forward) and Z(up) axes. For each particle, Integrate over area and time, to derive the center of air-mass momentum (p) before and after the passage of the airplane : dp/dt. (On a very calm morning, with no wind or turbulence, the center of air-mass and its momentum is stationary in Z(assume level un-accelerated flight), and equal to the True Airspeed in X pointing in the aft -X direction. Integrate over the area and you will find that the center and momentum of the particle and vector field has changed, with passage of the plane. This center of air mass and center of momentum will move forward(+X) and downward(-Z) relative to its original state. The equal and opposite momentum change with time dp/dt of the airplane is a force. We might label the -X component "drag" and the +Z component "lift"(careful: the airplane coordinate system is different from the stationary airmass). This is a dissipative system, so don't wait too long after the plane passes to record the vector field. We can observe this process in contrails on clear days when the high altitude air is cold and relatively moist. Sadly since we mostly view them from below with a projection along the Z, we miss the downward component of the momentum field. You can see this as a test pilot, flying as chase wing-man, in formation (projection in the Y-Z plane from behind or X-Z from the side). Expand this model to 3D to include lateral or Y axis flow and effects! I suggest this "p-dot"(dp/dt) of momentum-change explanation is better, than "pushing" or "pulling" the air downwards, because the later may confuse position and momentum in the view of the reader. This is also the first term(LHS) in the beautiful Euler-LaGrange equation, which would lead to an even more elegant analysis of this question!

As a new user, I will need to figure out how to attach the appropriate Figures and Equations to this post...-thanks

Note: The drag equation is really the ideal gas law, except density replaces m/V.

P/rho = R T :

Fluid interactions with solid bodies depend on the fluid properties and the geometry of the object. In the case of an aeroplane, we have air as our fluid and an aerofoil geometry. The aerofoil geometry is designed on purpose to force fluid under it preferentially to above it. This results in a pressure difference, which then leads to a buoyancy force that accelerates the wing according to Newton's second law (lift). Bernoulli's law is relevant for calculating the fluid problem.

So, to achieve flight, all you need is some well-designed aerofoils and some way of imparting an initial velocity. To keep flying you need to keep your speed high and to keep flying stably you need a well-designed aircraft with the centre of mass, centre of thrust and centre of lift being in the same position.

Essentially a fixed-wing aircraft flies because it moves through the air and has a fixed wing which is angled to the direction of airflow. A component of the drag force acting on the wing acts in the direction (up) opposite to the direction (down) of the aircraft's weight force.

An aeroplane wing acts like a weather vane responding to the relative flow of air. The basic effect can be obtained with a stiff, flat plate and a source of forward motion such as a propeller, gravity or launch momentum (e.g. children's paper planes). Refinements (such as aerofoil cross-sections) are introduced to mitigate the undesirable side-effects of flat plates (such as stalling).

No great argument with the other popular answers here but I will try to explain the basics of fixed wings in terms of Molecular Collisions. The following is rather a simplified explanation (ignoring things such as temperature, density, viscosity, compressibility, shear, boundary layers, turbulence, vortices, form drag, wing roughness, stiffness, skin friction, stalling, transmission by chain reactions, force couples etc).

A thought experiment. You sit at the bottom of a deep, water-filled swimming pool. You hold a table-tennis bat in one hand. Extend your arm and try to sweep the bat horizontally at constant speed through the water with the face of the bat firstly (a) vertical, then (b) horizontal, then (c) somewhere in between.

In case (a) the bat face is vertical and there will be the greatest resistance to forward motion. The resistance to forward motion can be explained by two broad effects.

The first effect is because the water molecules colliding with and rebounding elastically from the front face of the bat do so slightly faster and more frequently (on average) than the water molecules hitting the rear face of the bat. This is a simple consequence of the bat moving in the forward direction and the conservation of linear momentum in elastic collisions (think billiard balls hitting a large, massive, stiff, smooth, flat steel mirror). Each collision causes a change in the velocity of the bat. Because the frontal collisions are on average faster and more frequent than the rear collisions the net effect will be to reduce the forward velocity of the bat. In order to keep the bat moving at constant speed through the water you will need to expend muscular energy doing work against the resistance.

The second effect follows from the first effect. The molecules colliding with the front of the bat will be swept forwards causing an increase in pressure (a ram effect). This increase in pressure will act to further increase the air molecule velocities and rates of collision at the front face of the bat. The zone of increased pressure will grow in size ahead of the bat. Over time the continued growth of the high pressure zone will be offset by lateral diffusion of kinetic energy (high velocity molecules donating some of their velocity to surrounding slower-moving molecules by elastic collisions) and by mass flow of molecules past the edges of the bat to the lower pressure areas to the rear of the bat.

In case (b) the bat face is horizontal and the bat slides through the water with relatively little resistance.

In case (c) the bat face is tilted. The magnitude of the resistance depends on the angle of the bat face relative to the direction of motion. The resistance is greater when the bat face is near-vertical (steep angle of attack) compared to when the bat face is near-horizontal (shallow angle of attack). The resistance magnitude depends on the apparent cross-sectional area of the bat facing in the direction of motion. At shallower angle of attack fewer molecules impact the bat face, average angle of incidence of particles arriving at the bat face is greater causing reduced momentum exchange and there is less upstream pressure build up because it is easier (less obstruction) for molecules to escape the high pressure zone by flowing past the bat.

When the bat face is tilted upwards the net force on the bat is directed not backwards horizontally as in cases (a) and (b) but perpendicularly to the bat face (part backwards and part upwards). This can be explained by the geometry of molecular collisions at a flat surface moving through a stationary fluid.

A classical aerodynamicist might describe the face-perpendicular accelerations as combining components of both drag (backwards) and lift (upwards). If you tilt the bat so that the leading edge is tilted downwards then the net direction of the resistance to bat motion will be part backwards (drag) and part downwards ("negative lift"). Unqualified use of the term "lift" may cause confusion. It may be better to refer to components of wing-induced drag operating in specific directions (e.g. upwards, perpendicular to main airflow, perpendicular to wing surface, perpendicular to horizontal plane of the aircraft).

You can get a good feeling for the basic wing-induced drag effect by holding your hand, flat with fingers together, out of the window of an automobile when it is traveling fast (say 50 mph) and tilting your palm up and down and noting the forces which you feel when trying to keep your hand in the same position. (Probably best not to try a table tennis bat on public roads!).

Wings provide lift because they direct air downwards.

They direct air downwards in two ways. In part, the bottom of the wing slopes downward a bit and just pushes the air down as it moves forward through the air. But this is a small effect. The top of the wing is more important.

The top of the wing pulls the air down partially by providing a ramp. The rear portion of the top of the wing slopes down to a sharp trailing edge. The air, which is under pressure from the miles of air above it, follows that slope down the wing, and continues downward after the wing has passed.

But there is more to it than that. As the wing drives forward, the air that is deflected upward by the leading edge ends up being pinched between the layers of air above and the bulging top of the wing. That pinching makes the air speed up, not so differently from the way pinching a wet watermelon seed can send it flying. The inertia of the air that is farther from the wing forces the air that is closer to the wing to hug the wing's top surface, reaching the trailing edge much sooner than the corresponding molecules that headed along the bottom.

The asymmetry, of course, is key here. The bottom of the wing is more nearly parallel to the path of the air, with a bit of a downward slope all the way to the back, so it doesn't have the same pinching effect. (The asymmetry doesn't have to be in the shape of the wing. It can all be in the angle of attack. You are still creating a scenario where the air is pinched more on one side than the other.)

Of course there is no clear boundary between the layers of air that are doing the pinching and the air that is being pinched. But still, the force of the wing is felt most strongly by the air that is closest, and so that layer is accelerated the most. Each bit of air pinches the air below and is pinched against the air above, to a decreasing degree, until the effect is no longer noticeable quite some distance above the wing.

All this accelerated air is subject to the Bernoulli effect. Because it has been accelerated, its downward pressure on the wing is less than the upward pressure of the air below, and also the upward pressure on the air above is less than the ambient pressure. This causes even more air to move downward than otherwise would do so. Unless I am mistaken this is an important part of the downward deflection of the air.

The myth, then, is not that the Bernoulli effect is important. The myth is that there is an equal-time principle that is the reason the air atop the wing moves faster.

But the explanation is still incomplete because the Bernoulli principle itself is not obvious. The principle is often explained in terms of the low pressure causing the acceleration -- if you create an area of low pressure, air will indeed accelerate towards it. But if you blow into a tube with a construction, the decrease in pressure at the constriction will try to constrict it more. The upstream pressure from your lungs really is causing the decrease in pressure; it is not just the lower pressure that is causing the air to flow.

The way that increased pressure in your lungs can cause decreased pressure at the constriction is that your lungs give the air momentum. When the air finally leaves the tube that momentum is absorbed by the surrounding air, pushing it back like a mob pushing into a standing crowd. That momentum keeps some of the back pressure from being felt by the moving air in the tube. The higher the speed, the less momentum density, and the less back pressure.

In fact, in a steady state, inviscid, incompressible model, the question of what causes what becomes almost meaningless. The air speeds up because there is lower pressure in front, and there is lower pressure in front because of the velocity of the air. But in the case of an airplane, my understanding is that that the thrust of the engines is causing the acceleration of the air by more than just letting the downward sloping top of the wing recede from it. Even at high subsonic speeds where the air can no longer be treated as incompressible, the qualitative phenomenon that greater velocity leads to reduced pressure still applies. Calculating the effect just becomes more complicated.

Frequently, Bernoulli's principle is derived using conservation of energy along streamlines. I think my qualitative explanation using momentum is consistent with that.

The principle of lift is often explained using circulation. Again, I think that is just a different way of describing the same process. The different velocities along the top and the bottom constitute a net circulation.

Note: See "Why does the air flow faster over the top of an airfoil?" for additional answers to that part of the question of lift.

I'm late to the party here and I think the top vote-getters (Sklivvz, niboz) have adequately answered it, but I'll give my two cents anyway:

There are several ways to explain how an airplane flies. Some are more detailed than others, and unfortunately most popular explanation get it wrong. Here are some explanations that are useful, depending on the audience:

• The simplest explanation is that the wing pushes the air down and according to Newton's third law the air exerts an equal but opposite force up. The main way this happens is via the angle of attack, but the shape of the wing also plays a part. This suffices for most people, and should be the default explanation.

• A more detailed explanation would discuss the pressure difference between the two sides of the wing - since lift is a mechanical force it must be exerted on the surface of the wing and the only way air can do that is through pressure. So there must be a region of low pressure on the top of the wing and higher pressure on the bottom. Where does this come from? It comes from the air changing direction as it flows around the wing. Whenever air changes direction and follows a path that is curved there are pressure gradients with lower pressure on the inside of the curve.

• A still more detailed explanation would be to examine the Navier-Stakes equations and all the attendant math that goes with them. That's beyond the scope of this answer.

Holger Babinsky wrote a very readable paper called "How Do Wings Work?" that I'd recommend. It covers the middle answer quite well (and refutes a lot of the nonsensical explanation that are unfortunately all too common). Knowing a bit of calculus is helpful, but I think the article is readable without it. See http://iopscience.iop.org/0031-9120/38/6/001/pdf/pe3_6_001.pdf

Nib's answer is correct. The highly upvoted answer from Sklivvz starts promising but then throws in some incorrect statements:

Explanations showing a wing profile without an angle of attack are incorrect. Airplane wings are attached at an angle so they push the air down, and the airfoil shape lets them do so efficiently and in a stable configuration.

This incidence means that even when the airplane is at zero degrees, the wing is still at the 5 or 10 degree angle.

An asymmetric aerofoil creates lift at AoA zero. All fixed wing aircraft have asymmetrical aerofoils, only helicopters use symmetrical wing profiles in the rotor (due to these having no twisting moment). Fixed wing aircraft have wing twist: they have a positive angle of attack at the root, a negative AoA at the tip, and an average AoA as close to zero as possible, to minimise drag.

Indeed what makes the airplane fly is deflecting an airstream downwards. A flat plate can do this, and Bernoulli has no place in a flat plate. Subsonic aircraft don't use flat plates because they create a large amount of drag at angles of attack other than zero - in fact in a turbulent flow, even a flat plate at AoA zero creates more drag than a symmetrical wing profile such as NACA 0012.

The questioner continues objections because of other forms of flight that he points out. If we define flight only as a body creating lift using some way of moving clean air over an airfoil, then all the airfoil discussions are totally correct, and his examples are not relevant. If we loosen our definition of flight as getting a body off the ground for a sustain period beyond the effect of any initial ground-based propulsion, we still have balloons, rockets, and, to the point, many light aircraft with a thrust-to-weight ratio > 1, thereby allowing them to fly the aircraft stalled. The Harrier and the F-22 are prime examples, and the Osprey can be thrown in for a discussion of why helicopters fly.

In truth, all heavier-than-air flight is a combination of at least these two simple dynamics of airfoil lift and thrust energy surplus (that reserve available after satisfying forward motion for lift). And, of course, the whole calculus regarding wing lift gradients changes beyond the speed of sound and then at hypersonic speeds.

It is important to remember that a forward velocity is necessary for airfoil flight. That means, without some form of internal thrust, heavier-than-air airfoil flight is only a prolonged fall through the air. With any internal source of propulsion to sustain flight, we also give the pilot a way to create an energy surplus for maneuvering, increasing speed, or gaining altitude. Ask a pilot how he flies: "Angle-of-Attack, Air Speed, Altitude (repeat)". The airfoil is only a component.

Because of the obstruction of the wing, the air has to go around the wing, so the air pressure at the bottom of the wing is increased because the air at the bottom of the wing is compressed to go around the wing, and the air at the top of the wing is stretched around the wing, so the air pressure at the top of the wing decreases. So there's a pressure difference, and then there's a lift. Note: The bottom of the wing is windward, so the air is compressed, the pressure is high, and the top of the wing is leeward, so the air is stretched and the pressure is low. So lift cannot be explained by Bernoulli's theorem. Because Bernoulli's theorem does not consider the compression and stretching of fluid.

The following is a detailed explanation:

For example, at the top of the wing, the direction of air velocity at point A is the direction of the blue arrow. Because the blue arrow is inclined (note the angle between the blue arrow and the blue normal in the picture), the blue arrow tends to be far away from the wing along the normal direction at the top of the wing, so the air pressure at the top of the wing is stretched, so the air pressure at the top of the wing decreases, so there is a pressure difference (pressure gradient). This pressure difference changes the direction of air velocity, so the direction of air velocity at point B is the direction of the red arrow, and the red arrow is also inclined.... So the direction of air velocity will continue to change along the top of the wing. It should be noted that this pressure difference not only changes the velocity direction of the air on the top of the wing, but also generates the lift of the wing.

If there is no low pressure (negative pressure) at the top of the wing, will the airflow move downward? Obviously it won't move down. Wing lift comes from the low pressure at the top of the wing and the high pressure at the bottom of the wing. The downward movement of airflow is only the result of high and low pressure. Why is the top of the wing low pressure? Because the airflow tends to leave along the normal direction of the wing. Why is the bottom of the wing high? Because the airflow tends to approach along the normal direction of the wing.

The Newtonian explanation of flight based on the mass flow rate.

In stable cruise flight, wings with a positive angle of attack (AOA) fly through a mass of air each second (m/dt), and accelerate this air to a velocity (dv) downwards. This action crates a downward force (i.e. Force = ma = m/dt x dv). The reaction generates an equal and opposite upward force that provides lift. Lift is the vertical component of the upward force. Simply put, when the air goes down and the airplane goes up.