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• the local Reynolds number,
• the pressure gradient,
• wing sweep and
• disturbances like bugs, rivet heads or turbulators.

The "rooftop" distribution of the 6-digit NACA airfoils did help, though, because it gives them a higher critical Mach number than the "peaky" distributions of earlier airfoils. The suction peak near the nose of older airfoils will lead to local supersonic flow at a lower flight Mach number, and increased drag from the shocks which would follow. Most important for its low drag, however, was the very smooth wing surface of the P-51 with no gaps ahead of the spar. See this rec.aviation.military post for details.

• the local Reynolds number,
• the pressure gradient,
• wing sweep and
• disturbances like bugs, rivet heads, waviness or turbulators.

The "rooftop" distribution of the 6-digit NACA airfoils did help, though, because it gives them a higher critical Mach number than the "peaky" distributions of earlier airfoils. The suction peak near the nose of older airfoils will lead to local supersonic flow at a lower flight Mach number, and increased drag from the shocks which would follow. Most important for its low drag, however, was the very smooth wing surface of the P-51 with no gaps ahead of the spar. See this rec.aviation.military post for details.

Measurements on the P-63

The Bell P-63 Kingcobra used early laminar airfoils, the NACA 66(215)-116 at the root and the NACA 66(215)-21 at the tip. British testing of its wing suggests that the build quality of the period's metal wings was insufficient to maintain laminar flow. From the linked Wikipedia article:

The RAE first tested it in an "as delivered" configuration. The wing airfoil was designed to support laminar flow to 60% of chord. In the "as delivered" configuration, a profile drag was measured which was representative of the wing section with boundary layer transition at the leading edge (0% laminar flow). Reducing the surface roughness reduced the drag at low lift coefficients to a level representative of laminar flow to 35% of chord. Measurements were made of the surface waviness. This showed peak wave amplitudes, above the mean, of approximately 0.011 inches (0.28 mm) over a two-inch (5.1 cm) span. The standard waviness criteria shows the critical wave height to be 0.0053 inches (0.13 mm) for this application. To reduce the waviness, RAE personnel stripped the wing to bare metal. The wing was then sprayed with two coats of primer paint and a coat of paint type filler. After the paint was dry, it was sanded in a chordwise direction, using sanding blocks, whose curvature matched the local surface curvature. This was repeated several times. Surface waviness was then measured and found to be no more than 0.005 inches (0.13 mm). In flight, this configuration was found to have a profile drag representative of boundary layer transition at 60% of chord. This gave researchers an idea of what level of wing surface quality was required to actually get the benefits of laminar flow airfoils.

The "rooftop" distribution of the 6-digit NACA airfoils did help, though, because it gives them a higher critical Mach number than the "peaky" distributions of earlier airfoils. The suction peak near the nose of older airfoils will lead to local supersonic flow at a lower flight Mach number, and increased drag from the shocks which would follow. Most important for its low drag, however, was the very smooth wing surface of the P-51 with no gaps ahead of the spar. See this article for details.

The "rooftop" distribution of the 6-digit NACA airfoils did help, though, because it gives them a higher critical Mach number than the "peaky" distributions of earlier airfoils. The suction peak near the nose of older airfoils will lead to local supersonic flow at a lower flight Mach number, and increased drag from the shocks which would follow. Most important for its low drag, however, was the very smooth wing surface of the P-51 with no gaps ahead of the spar. See this rec.aviation.military post for details.

In still air every boundary layer starts laminar. How soon it transitions to a turbulent boundary layer depends on:

• the local Reynolds number,
• the pressure gradient,
• wing sweep and
• disturbances like bugs, rivet heads or turbulators.

Wing sweep will also make it hard to maintain laminar flow. As explained here, on a swept wing only the speed component perpendicular to the wing will be affected by it, so the accelerating flow past the stagnation point will curve inwards on a sweptback wing. At the same time, viscosity will slow down the flow near the wing skin. The consequence is a twist in the speed distribution over the boundary layer, which destabilizes the laminar flow and leads to early transition.

A C-172 with its four-digit NACA airfoil has the peaky upper surface which will trip the boundary layer very early on the upper surface. On the lower surface laminar flow will last a little longer but will be destabilised by gaps in the surface, so most of the flow on the C-172 is turbulent. On an airliner the Reynolds number is in the tens of millions, so transition will be very early and very little laminar fraction remains. This is mostly found near unswept leading edges like the engine nacelles. Only with advanced technologies like boundary layer suction will it be conceivable that a larger part of an airliner wing can be kept laminar.

Laminar separation sometimes happens when the flow separates shortly after negotiating the nose, like on a five-digit NACA airfoil or an undimpled golf ball. This leads to an abrupt stall and should best be avoided. Normally, the boundary layer transitions into the turbulent state and stays attached until the turbulent boundary layer separates, either at the trailing edge, or progressively further ahead of it when the airfoil stalls.

In still air every boundary layer starts laminar. How soon it transitions to a turbulent boundary layer depends on:

• the local Reynolds number,
• the pressure gradient,
• wing sweep and
• disturbances like bugs, rivet heads or turbulators.

Wing sweep will also make it hard to maintain laminar flow. As explained here, on a swept wing only the speed component perpendicular to the wing will be affected by it, so the accelerating flow past the stagnation point will curve inwards on a sweptback wing. At the same time, viscosity will slow down the flow near the wing skin. The consequence is a twist in the speed distribution over the boundary layer, which destabilizes the laminar flow and leads to early transition.

A C-172 with its four-digit NACA airfoil has the peaky upper surface which will trip the boundary layer very early on the upper surface. On the lower surface laminar flow will last a little longer but will be destabilised by gaps in the surface, so most of the flow on the C-172 is turbulent. On an airliner the Reynolds number is in the tens of millions, so transition will be very early and very little laminar fraction remains. This is mostly found near unswept leading edges like the engine nacelles. Only with advanced technologies like boundary layer suction will it be conceivable that a larger part of an airliner wing can be kept laminar.

Laminar separation sometimes happens when the flow separates shortly after negotiating the nose, like on a five-digit NACA airfoil or an undimpled golf ball. This leads to an abrupt stall and should best be avoided. Normally, the boundary layer transitions into the turbulent state and stays attached until the turbulent boundary layer separates, either at the trailing edge, or progressively further ahead of it when the airfoil stalls.