incarcerated |
04-03-2011 18:23 |
1 Attachment(s)
Quote:
Originally Posted by Buffalobob
(Post 384906)
Bullet shape is important. Shallow boat tail angle and no meplat or very small meplat, and short bearing surface seem to be important.
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Ahhhh…
I can see the meplat being an issue, as a hollow point degrades the BC of any rifle bullet, but I had not considered the boat tail. If the instability derives from changes in the bullet’s shockwave as it crosses the transonic plane, which was my uneducated guess (and which turns out to be not quite accurate), my attention would be on the front of the bullet. Being a short-range shooter (nothing much past 200) and fond of flat base bullets, I would not have considered the effect of the vortex following the bullet. The short bearing surface will help at both ends of the bullet, or put differently, will result from reducing bluntness up front and lengthening the taper of the boat tail.
Will an epoxy filled nose and a rearward center of gravity help things here?
And where is Longrange1947? I’d like to hear his input on this.
A little Google-fu helped throw some light on the issue for me: rather than looking at the bullet, it got me looking at aerodynamics, specifically, the transonic zone:
Quote:
http://answers.yahoo.com/question/in...3211733AA9JEeY
The Transonic zone is a condition of speed when some parts of the air flowing over an object, such as an airfoil are supersonic already and some parts are not. The term to watch between the transonic zone and the supersonic zone is called the Critical Mach number...where the airlow in some parts of an object, again, the typical example would be an airfoil..reaches the speed of sound, even though other parts of the aircraft has not crashed through the sound barrier. When this condition is reached, it will create a weak shock wave.
A shock wave is a condition where a there is always a rapid rise in pressure, density and temperature. As the aircraft approaches the transonic zone--Mach .80 to Mach 1.2--the pressure waves do not have time to move out of the way of the oncoming aircraft since it is traveling along with them. The waves compress and the air becomes far more dense. When the plane meets the compacted air, it hits with a jolt and a series of shock waves builds up perpendicular to the direction of flight. The first shock wave attaches itself to the center of the wing's upper surface as the airflow there reaches Mach 1, As the plane's speed increases, the air under the wing also reaches mach 1 and a second shock wave forms....
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Quote:
http://www.centennialofflight.gov/es..._Flow/TH19.htm
An airplane flying well below the speed of sound creates a disturbance in the air and sends out pressure pulses in all directions. Air ahead of the airplane receives these "messages" before the airplane arrives and the flow separates around the airplane. But as the plane approaches the speed of sound, the pressure pulses merge closer and closer together in front of the airplane and little time elapses between the time the air gets a warning of the plane's approach and the plane's actual arrival time. At the speed of sound, the pressure pulses move at the same speed as the plane. They merge ahead of the airplane into a "shock wave" that is an almost instantaneous line of change in pressure, temperature, and density. The air has no warning of the approach of the airplane and abruptly passes through the shock system. There is a tendency for the air to break away from the airplane and not flow smoothly about it; as a result, there is a change in the aerodynamic forces from those experienced at low incompressible flow speeds….
At subsonic speeds, drag was composed of three main components—skin-friction drag, pressure drag, and induced drag (or drag due to lift). At transonic and supersonic speeds, there is a substantial increase in the total drag of the airplane due to fundamental changes in the pressure distribution.
This drag increase encountered at these high speeds is called wave drag. The drag of the airplane wing, or for that matter, any part of the airplane rises sharply, and large increases in thrust are necessary to obtain further increases in speed. This wave drag is due to the unstable formation of shock waves that transforms a considerable part of the available propulsive energy into heat, and to the induced separation of the flow from the airplane surfaces. Throughout the transonic range, the drag coefficient of the airplane is greater than in the supersonic range because of the erratic shock formation and general flow instabilities. Once a supersonic flow has been established, however, the flow stabilizes and the drag coefficient is reduced…
It is a large loss in propulsive energy due to the formation of shocks that causes wave drag. Up to a free-stream Mach number of about 0.7 to 0.8, compressibility effects have only minor effects on the flow pattern and drag. The flow is subsonic everywhere. As the flow must speed up as it proceeds about the airfoil, the local Mach number at the airfoil surface will be higher than the free-stream Mach number. There eventually occurs a free-stream Mach number called the critical Mach number at which a supersonic point appears somewhere on the airfoil surface, usually near the point of maximum thickness, and indicates that the flow at that point has reached Mach 1. As the free-stream Mach number is increased beyond the critical Mach number and approaches Mach 1, larger and larger regions of supersonic flow appear on the airfoil surface. In order for this supersonic flow to return to subsonic flow, it must pass through a shock (pressure discontinuity). This loss of velocity is accompanied by an increase in temperature, that is, a production of heat. This heat represents an expenditure of propulsive energy that may be presented as wave drag. These shocks appear anywhere on the airplane (wing, fuselage, engine nacelles, etc.) where, due to curvature and thickness, the localized Mach number exceeds 1.0 and the airflow must decelerate below the speed of sound. For transonic flow, the wave drag increase is greater than would be estimated from a loss of energy through the shock. In fact, the shock wave interacts with the boundary layer so that a separation of the boundary layer occurs immediately behind the shock. This condition accounts for a large increase in drag that is known as shock-induced (boundary-layer) separation….
At a free-stream Mach number greater than 1, a bow shock appears around the airfoil nose. Most of the airfoil is in supersonic flow. The flow begins to realign itself parallel to the body surface and stabilize, and the shock-induced separation is reduced.
This condition results in lower drag coefficients. Supersonic flow is better behaved than transonic flow and there are adequate theories that can predict the aerodynamic forces and moments present. Often, in transonic flow, the flow is unsteady, and the shock waves on the body surface may jump back and forth along the surface, thus disrupting and separating the flow over the wing surface. This sends pulsing, unsteady flow back to the tail surfaces of the airplane. The result is that the pilot feels a buffeting and vibration of both wing and tail controls. This condition occurred especially in the first airplane types to probe the sound barrier....
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