# Why does supersonic flight detach airflow from a wing?

I've read in several answers to questions that when a wing passes the speed of sound the airflow will become detached from the craft towards the rear quarter of the wing (thus making things like elevators ineffective.)

Why does this happen at Mach Speed? Can it happen at lower speeds as well (in straight and level flight, ie, not stalled)?

Flow separation happens when the pressure gradient of the airflow along the flow path becomes too steep. In subsonic flow, the oncoming air is first decelerated ahead of the wing, then swiftly accelerated when it flows around the strongly curved nose section of a wing. This acceleration is the consequence of the wing's curvature.

After this acceleration, how will the airflow behave? Will the airflow keep moving in a straight line without further acceleration or deceleration? (which implies that airflow will separate from the wing surface, creating a local vacuum) Or will it faithfully stick to the wing surface? In reality, the air settles at a compromise between the straight path and following the contour, creating decreasing pressure along a surface with increasing curvature and increasing pressure along surfaces with decreasing curvature. More precisely, it is always in an equilibrium between inertial, viscous and pressure forces.

This suction not only bends the airflow into following the wing's contour, but also accelerates the air ahead of it. The lower the pressure, the more the air speeds up, such that the total energy of air (the sum of pressure and kinetic energy) stays constant. Therefore, pressure and local speed change in sync.

When the curvature decreases further downstream, the flow path becomes straighter and pressure rises again. However, the air particles close to the wing slow down because of friction. The layer of air where this slowing is noticeable is called boundary layer. In it, the deceleration effects due to pressure rise and due to friction add up, and at some point the air comes to a standstill relative to the wing. Where this happens, static air will collect and build up, causing flow separation. Thankfully, the exchange of air across a turbulent boundary layer kicks the slowest air particles downstream, so at moderate angles of attack the air still moves along until it reaches the trailing edge. Only when the suction peak around the nose becomes very high at high angle of attack, the consequent steep pressure rise along the remaining flow path overpowers the possibilities of the turbulent boundary layer, the air decelerates completely and the flow separates. This is a wholly subsonic affair.

If the wing moves at high subsonic speed, the curvature-created suction accelerates the flow such that it reaches supersonic speed. Now something odd happens: Supersonic flow accelerates further when subsonic flow would decelerate. This is caused by the change in density which is dominant at supersonic speed. Incompressible (= very slow) flow has constant density, and all speed changes affect pressure. At Mach 1, the pressure and density changes are of equal magnitude, and in supersonic flow the density changes dominate. Now we have a supersonic pocket of air on the upper surface of the wing where speed increases and density decreases downstream, and the surrounding subsonic air sees little change in density. This picture should give you some idea how it looks: ![Lambda shock in high subsonic flow][1]

The whole wing moves at Mach 0.68. Compare the green color at some distance from the wing with the scale on the left side, which gives the Mach number for each color shade. At the airfoil nose, you see a blue area. This is where the air decelerates - it gets pushed together by the approaching wing. Now follow the colors along the upper side - they quickly turn green, yellow and red as the air is accelerated into the low pressure area (remember, low pressure equals high speed, so the reddest area has the highest local flow speed and the lowest pressure). In subsonic flow, the suction peak would be somewhere between 20% and 30% of chord, and the colors would slowly change back to yellow and green if you move further downstream. Now we have local supersonic flow (everything redder than light orange is supersonic here), and instead of slowly decelerating, the air is speeding up to a maximum Mach number of 1.23 at almost 60% of chord length.

This cannot last, and at some point this supersonic pocket collapses. This happens instantly in a shock, and as you know, in a straight shock density increases suddenly and speed decreases such that the Mach number after the shock is the inverse of the Mach number ahead of the shock. In the picture above, boundary layer effects create a lambda shock, which has its name from the Greek letter which looks like the shock pattern here. Aft of the shock, you have subsonic flow again, and a much thicker boundary layer which moves very slowly (blue shade). This is due to the energy conversion through the shock, which converts kinetic energy into heat. But the flow is still attached - even this shock did not cause separation.

If this pressure rise is big enough, the boundary layer will come to an instantaneous standstill, and then the flow separates. This is the shock induced separation you asked about. Unfortunately, the picture above is the best I could find, and I have none with separated flow aft of the shock. But it helps to show that the center of pressure moves aft. This causes a strong nose-down moment. Also, with higher subsonic Mach numbers directional stability decreases. Now even more nasty things can happen: The location of the shock might move forward and back. This changes the size of the supersonic area, causing lift changes. On a horizontal tail, this will also cause pitch changes. If you change elevator position slightly with such a shock on the horizontal tail, the lift change could be severe and in the opposite direction of what you would expect. This causes total loss of control, just when you need the control surfaces to counteract the Mach effects mentioned above. Also, the shock position could oscillate, causing a buzzing sound and, if you are really unlucky, coupling into an elastic eigenfrequency of your structure, resulting in flutter. Not only on the tail surfaces, but also on the wing, affecting the ailerons as well. Now you can start to see what scared the early pioneers about flying near Mach 1 and why they spoke of a "sound barrier".

If you fly fully supersonic, this effect goes away because now the shock moves to the trailing edge and stays there. Now everything will be calm again because the shock location stays fixed. This effect was first experienced and survived on April 9, 1945 by [Hans Guido Mutke in a Me-262][3], which briefly flew fully supersonic in a dive. However, even in fully supersonic flow separation is possible, but then because the air flow will not bend more than what can be caused by total vacuum. In hypersonic flow the density changes become so severe that pockets of "air" are possible which do not contain any air, but a vacuum. But this is more an academic case, except for reentry vehicles with a blunt, rear-facing base.

• Ok. I'm a professional software engineer by-trade (and pulling no punches, a damn-fine one), and I have to say, I'm overtly jealous of the vernacular of aeronautical engineers. You guys have the coolest damn terms. I could read your stuff for hours (and, shockingly, understand most of it, but that I attribute to the physics and math training of long-ago). Nice writeup. – WhozCraig May 24 '14 at 4:25
• @WhozCraig: Thank you for the really nice comment! In order to improve the post, could you please point out what is not easy to understand? I would like that readers understand all of it, and I appreciate your help! Thankfully, Stack Exchange allows to edit posts, so I can make it better. – Peter Kämpf May 24 '14 at 6:16
• The fourth paragraph took a few passes before it started to finally gel. The rest was cake to sink in, but that paragraph took some doing. I know its a hard subject to describe, particularly to people that are foreign to most of even the basic concepts, but you did pretty darn well, I have to say. If you changed anything, it would be honing that one paragraph to the less educated masses, but I'm not convinced that would be your target audience to begin with. – WhozCraig May 24 '14 at 6:55
• I have to admit that I have no easy and intuitive explanation for the differences between subsonic and supersonic flow. It took me a while to explain pressure around an airfoil without vortices and all what comes with the totally unintuitive potential flow theory. I am still working on understanding supersonic flow myself. – Peter Kämpf May 24 '14 at 8:58
• Just curious to know which software did you use for the simulation and more importantly which assumptions and computations types have been used (Euler, RANS, turbulence model and so on)? Explanations are really clear to me, nice job (even though I do have some background in that field which might help to grasp the thing)! – Ludovic C. May 24 '14 at 10:03

Why does this happen at Mach Speed? ...can it happen at lower speeds?

It can also happen at lower speeds, it depends on how the wing is designed and the characteristics of the airfoil (camber thickness profile).

At transonic speeds (0.7 - 1.0 Mach) you can have portions of the airfoils in a supersonic region, meaning that you will have a shock front over (and maybe under) your wing. If the shockwave is strong enough, the flow behind it will be (partially) separated.

When the aircraft reaches Mach 1 the presence of a shock is guaranteed.

Image from wiki

EDIT

In the comments more questions have arisen, I will try to address them.

why does the Shockwave appear?

Short answer: To go back from supersonic to subsonic flow. Supersonic flow is hard to decelerate without a shock, since the air molecules don't "know" what is ahead. The speed of sound is also the speed of small pressure changes, so any signals of what is coming will not reach the air ahead of the shock wave. The air streams along, blissfully unaware of what is coming, until things cannot be maintained and change with a bang.

When the shock wave is so extended to reach the ground, it is called Sonic boom: see the section Sonic boom and sound barrier

Because of friction, the air molecules next to the body have no velocity relative to the bodies surface. The molecules a bit further away will be able to move, but because of friction with the molecules attached to the body, they will be slowed down. This phenomena is called a boundary layer. In normal conditions this is the way roughly half of the profile drag is created. The other half is pressure drag. In separated flow, friction drag disappears, but since the separated air mass is at a lower pressure than static pressure, and because it sits on the backward-facing part of the airfoil, its pressure drag contribution is massive.

why would the shockwave create turbulence?

Because is an anisotropic event, it is chaotic in nature, it increases the entropy of the air molecules.

Temperature, density, pressure and velocity change by such large amounts at the shock wave (depending on the speed of the supersonic flow, with exact quantities given by solving the Rankine–Hugoniot equations) and in such an infinitesimally small space that the flow downstream becomes extremely chaotic and non-laminar.

Can it happen at lower speeds as well (in straight and level flight, ie, not stalled)?

A shock wave can only happen if you have supersonic flow over the wings. Flow separation happens at all speeds. In the end, all flow will separate at the trailing edge.

It happen on mach speed because that is when the wing will outspeed the pressure wave of it trying to cut through the air resulting in a shockwave when the air slows down again to subsonic speeds this shock is called a recompression shock.

The shockwave is the cause of the separation. And as you see it can happen at lower than 1 Mach. The minimal airspeed at which supersonic flow happen is called the critical mach number. But the speed at which the drag cause by the shock wave become significant is the drag divergence mach number.