WWII A/C: Maximum Mach Number & Airspeed in Dives

[Zipper730]

Basically I'm interested in the maximum mach number of the various A/C that were either conceived in, or served in WWII.

My interest includes the maximum mach number ever safely attempted in tests and/or combat, and (if possible) the maximum the plane would take before it either became uncontrollable or structurally disintegrated.

There are some designs which I know about in varying degrees, others very little to nothing and I'm interested in finding out the dive performance of the following

1. Supermarine Spitfire (Mk.I-VI): The early Spitfires had more elasticity in the wings, which limited the maximum dive speed. I know by the Mk.VII or VIII, they managed to redesign the wings to allow an Mmo of Mach 0.85. It was possible to go beyond that speed, and Mach 0.89 was achieved in one case.
   
2. Supermarine Spitfire (Late Merlin & Griffon Variants): I remember at least two PR variants that achieved remarkable high-speeds in dives. The first was used to test the concept of a stabilator for the M.52, and they achieved Mach 0.92, and there was a PR.XIX, I think, that achieved Mach 0.94 after going out of control from high altitude during atmospheric testing. I remember some Spitfires were fitted with a modified leading-edge that served to reduce the effective camber of the wing: My interest is mostly if these planes were fitted with them.
3. Chance-Vought F4U: I've been told it was the first US carrier-based fighter to be vulnerable to mach-tuck, but I don't know exactly when control force heaviness began due to either airspeed or mach, the Mmo, and the absolute max was.
4. Curtiss SB2C: Though it used dive-brakes normally, the aircraft if put into a dive without them could get fast enough to suffer compressibility effects. I'm curious as to how fast they could go?
5. De Havilland Mosquito: If I recall it's maximum IAS was 430 mph+, but I don't know what the Mmo was.
   
6. Republic P-47C/D: I remember hearing that the plane's maximum tactical mach number was 0.72, other sources said control problems started as early as 0.67-0.69: I'm not sure which is correct, and what the Mmo and absolute maximum were. I know some P-47's were fitted with a type of dive-recovery flap, and I'm curious how much they increased the maximum mach number achievable.
   
7. Republic P-47N: It had a different wing than the earlier P-47, not just in terms of a clipped wingtip, but the basic cross-section as I understand it was altered as well.
8. Grumman F6F: Their top-speeds were a little slower than the F4U, but I'm curious how it compared to the F4U in terms of Mmo
9. Boeing B-29: The a/c's maximum dive speed was 300 mph indicated and while I can compute indicated airspeed to true airspeed, and derive mach number: I'm not sure how long the plane would have to dive from service/absolute ceiling to reach that speed.

Weird questions I know

 

[drgondog]

The simple answer is that no WWII airplane had a "design' maximum Diving Speed. All aircraft had a design Limit and Ultimate load of 1.5xLimit Stress based on the material limits of the airframe structure but too little was known about high speeds approaching M=1.

Next Point - The airplane (usually wing) has three important velocity regimes. The first is compressible flow but below any local M=1 values. The second is the Critical Mach transition in which local bubbles between leading edge and maximum thickness to chord ratio approach supersonic and shock waves form and dissipate, until finally sustained supersonic flow (M=1 at Mcr) creates and sustains a shock wave. More of this below.

The aerodynamic loads and pressure distributions on the airframe (mostly focused on Wing and Empennage) were passed on to airframe structures teams so that the airframe design components (spars, ribs, fittings, bolts, rivets, skin, longerons, etc) could be individually examined for the maximum stress due to the 'worst flight condition pressure distribution over the wing and tail surfaces, as well as Torsion and Tension and compression loads to members designed to carry the loads.

True, that dive conditions during the dive (before pullout at which point the applied loads are assumed due to angle of attack/lift loads) generates Dynamic Pressure loads different from Lifting loads - which applies hundreds of pounds per square inch to the surfaces normal to the free stream. That said there were no good analytical methods to calculate them and flight testing had to be conducted to evaluate the behavior as well as to note observable issues (skin buckling, popped rivets, bent spars, etc) as the airframe stresses went past Limit stress on way to Ultimate Stress where the airframe no longer bent - but failed, locally, until a major structures component failed completely.

When completed and analyzed V-n diagrams which showed positive and negative N (G) as function of banked turn/pullout loading also displayed the Q (Dynamic Pressure) limit for top speed. That top speed threshold translated to the dive speed Placard for operational limit. This velocity 'point' was sub Critical Mach but usually in a range from .62M to .75M but always below Mcr (Critcal Mach) at which point the local airflow over the wing reached M=1 LOCALLY. For conventional wings of common T/C (14-16% like P-38, FW 190, Bf 109, P-47, F4-U, F6F), the shock wave developed at M=1 locally usually did tow things (both 'bad') by moving the Center of Pressure over the wing from closer to 25% Chord to closer to 50% Chord and the resulting boundary layer separation resulted in blanking out the Elevator. Both effects tended to create a Pitch Down effect.

For a Laminar flow (Mustang NAA/NACA 45-100 for example) wing, the CP moved very little from the Max T/C point of 45%. Separation did occur at that point resulting in some buffeting but the Mustang did not pitch down as the Moment Coefficient did not change dramatically due to the CP migration during the velocity gradient approaching M=1.

The Spitfire had by far the best T/C ratio beneficial to a lower velocity gradient from the leading edge to the 25% T/C (maximum thickness of airfoil for the Spit and FW 190 and Bf 109, etc) so it achieved a higher dive speed before the 'Tuck' effect occurred.

The Tuck effect introduced the requirement for positive elevator defection which in turn imposed higher 'angle of attack' loading at very high speed - sometimes resulting in yanking an empennage and nearly always resulting in shear panel buckling loads as main spars deflected past design limits analyzed previously.

Summary - your question implies predictable dive limits based on Mach number. Too many variables.

 

[Zipper730]

The simple answer is that no WWII airplane had a "design' maximum Diving Speed. All aircraft had a design Limit and Ultimate load of 1.5xLimit Stress based on the material limits of the airframe structure

That depends on nation, the United States had an ultimate load of 1.5 x design limit; the Germans used an ultimate load of 1.7 x design limit.

Next Point - The airplane (usually wing) has three important velocity regimes. The first is compressible flow but below any local M=1 values. The second is the Critical Mach transition in which local bubbles between leading edge and maximum thickness to chord ratio approach supersonic and shock waves form and dissipate, until finally sustained supersonic flow (M=1 at Mcr) creates and sustains a shock wave. More of this below.

The supersonic areas would be called supercritical, right?

When completed and analyzed V-n diagrams which showed positive and negative N (G) as function of banked turn/pullout loading also displayed the Q (Dynamic Pressure) limit for top speed.

V-n = Vne?

That top speed threshold translated to the dive speed Placard for operational limit. This velocity 'point' was sub Critical Mach but usually in a range from .62M to .75M but always below Mcr (Critcal Mach) at which point the local airflow over the wing reached M=1 LOCALLY.

I thought the Spitfire, P-51, and Me-262 all had some super-critical airflow on them as they were all capable of reaching/exceeding Mach 0.8.

For conventional wings of common T/C (14-16% like P-38, FW 190, Bf 109, P-47, F4-U, F6F), the shock wave developed at M=1 locally usually did tow things (both 'bad') by moving the Center of Pressure over the wing from closer to 25% Chord to closer to 50% Chord and the resulting boundary layer separation resulted in blanking out the Elevator. Both effects tended to create a Pitch Down effect.

I thought the problem was the interference of airflow over the tail and the loss of downwash. I didn't think there was much shift in the center of pressure until you were nearly at Mach 1 (upwash disappears), though the pitch down would be equivalent to such a shift.

For a Laminar flow (Mustang NAA/NACA 45-100 for example) wing, the CP moved very little from the Max T/C point of 45%. Separation did occur at that point resulting in some buffeting but the Mustang did not pitch down as the Moment Coefficient did not change dramatically due to the CP migration during the velocity gradient approaching M=1.

I figured it had to do with the thickest part of the wing located further aft producing an effect of a thinner wing (t/c is around 16%).

The Fw-190's dive speed was limited early on by the either the rudder or elevator control balance or something, later on though they could dive with or faster than the P-51.

The Spitfire had by far the best T/C ratio beneficial to a lower velocity gradient from the leading edge to the 25% T/C (maximum thickness of airfoil for the Spit and FW 190 and Bf 109, etc) so it achieved a higher dive speed before the 'Tuck' effect occurred.

The lower the velocity gradient means a higher Mcr?

The Tuck effect introduced the requirement for positive elevator defection which in turn imposed higher 'angle of attack' loading at very high speed - sometimes resulting in yanking an empennage and nearly always resulting in shear panel buckling loads as main spars deflected past design limits analyzed previously.

I though the higher AoA was caused by the L/D dropping, and thus requiring more lift to keep the plane flying.

Summary - your question implies predictable dive limits based on Mach number. Too many variables.

The aircraft usually did have maximum TAS and IAS, and those if they correlated to an altitude can yield a mach number. For example 350 knots at 30,000 feet is Mach 1

 

[drgondog]

That depends on nation, the United States had an ultimate load of 1.5 x design limit; the Germans used an ultimate load of 1.7 x design limit.

The supersonic areas would be called supercritical, right? No. Supercritical refers to low drag airfoil (frequently with max T/C near trailing edge (i.e 75-80%) to drive Drag Divergence/Mcr velocity much closer to M=1.0

V-n = Vne?Yes

I thought the Spitfire, P-51, and Me-262 all had some super-critical airflow on them as they were all capable of reaching/exceeding Mach 0.8. Probably all WWII were capable of Mcr locally but all had slightly different to greatly different Drag Divergence/Mcr velocities

I thought the problem was the interference of airflow over the tail and the loss of downwash. One of the problems. I didn't think there was much shift in the center of pressure until you were nearly at Mach 1 (upwash disappears), though the pitch down would be equivalent to such a shift. The CMac is airfoil related in this case - not the immersion of horizontal stabilizer/elevator. The addition of wing fence/brake was designed to solve this issue.

I figured it had to do with the thickest part of the wing located further aft producing an effect of a thinner wing (t/c is around 16%). Both the lower velocity gradient and the absolute velocity was enhanced by the location of the max T/C at 45%

The Fw-190's dive speed was limited early on by the either the rudder or elevator control balance or something, later on though they could dive with or faster than the P-51.
Source?

The lower the velocity gradient means a higher Mcr? All other features being equal - yes

I though the higher AoA was caused by the L/D dropping, and thus requiring more lift to keep the plane flying.

Two discussions - first is dive pull out when out of the severe compressibility effects discussed above - that is elevator discussion and nothing to do with L/D, but everything about Angle of Attack and the aforementioned AoA induced G loads. The Tuck is caused by an aft movement of CP causing a much more negative CMac than at zero lift - and as an example the P-38 didn't have elevator control authority while immersed in the Mcr induced flow separation.

The aircraft usually did have maximum TAS and IAS, and those if they correlated to an altitude can yield a mach number. For example 350 knots at 30,000 feet is Mach 1

True - and the placard references Limit dive speeds as a function of IAS/Altitude so the pilot doesn't have to do the conversion. The structural discussion is better reduced to Dynamic Pressure encountered.

OTOH M=1 at 30,000 feet is 589 kts and 678 mph. The easy calc is multiply 29.04*sqrt (T) with T in Rankin. For STP at SL, T=411.8 Rankin at 30,000 feet.

 

[drgondog]

That depends on nation, the United States had an ultimate load of 1.5 x design limit; the Germans used an ultimate load of 1.7 x design limit.

Interesting - Source for German standards? Aluminum has a Failure point at approximately 1.5Yield. What made German aluminum so robust past Yield?

 

[pbehn]

Interesting - Source for German standards? Aluminum has a Failure point at approximately 1.5Yield. What made German aluminum so robust past Yield?

Possible depending on what the design limit is as a percentage of yield.

 

[pbehn]

Didnt the propellor fall off one of the spitfires used in the high speed diving test? I cant believe any aircraft with propellor on the front got anywhere near the sound barrier

 

[wuzak]

Interesting - Source for German standards? Aluminum has a Failure point at approximately 1.5Yield. What made German aluminum so robust past Yield?

Its a safety factor Bill. So for the design load the Germans were calculating stresses using 13% greater than in teh US, meaning that the structure has to be stronger/heavier.

I would like to know where that information came from too.

 

[wuzak]

Didnt the propellor fall off one of the spitfires used in the high speed diving test? I cant believe any aircraft with propellor on the front got anywhere near the sound barrier

Yes, that is correct. The reduction gear failed and the propeller came adrift.

On another occasion the supercharger exploded.

But the high mach number (0.9M) had been achieved in an earlier test without damage to the aircraft.

 

[pbehn]

Yes, that is correct. The reduction gear failed and the propeller came adrift.

On another occasion the supercharger exploded.

But the high mach number (0.9M) had been achieved in an earlier test without damage to the aircraft.

Two fails and one pass doesnt get you into regular service, I would say that it shows that the spitfire was not safe at 0.9M

 

[drgondog]

Possible depending on what the design limit is as a percentage of yield.

Its a safety factor Bill. So for the design load the Germans were calculating stresses using 13% greater than in teh US, meaning that the structure has to be stronger/heavier.

I would like to know where that information came from too.

I am aware that the 1.5 multiple of the Design Stress is a 50% safety factor. The airframe biz has standard tables for Tensile strength design allowable for each material based on industry wide strength of Material standards, While I mention Yield Point as the approximate design Limit tensile strength, the 'book values' are the industry airframe standards. For most aluminums and steels, 1.5xYield is the threshold for Failure point.

!,7 factor over US industry standard Tensile strength implies one of two things. 1.) They (Germany) had a permissible Tensile strength significantly below Yield but near same for Failure, or 2.) same Yield but based on milling/extrusion/forming improvements but more impressive ductility curve to get a higher stress Failure point at 70% greater than their standard design limit which is same as US.

I only use Yield in a discussion as they are findable more easily.

 

[pbehn]

I am aware that the 1.5 multiple of the Design Stress is a 50% safety factor. The airframe biz has standard tables for Tensile strength design allowable for each material based on industry wide strength of Material standards, While I mention Yield Point as the approximate design Limit tensile strength, the 'book values' are the industry airframe standards. For most aluminums and steels, 1.5xYield is the threshold for Failure point.

!,7 factor over US industry standard Tensile strength implies one of two things. 1.) They (Germany) had a permissible Tensile strength significantly below Yield but near same for Failure, or 2.) same Yield but based on milling/extrusion/forming improvements but more impressive ductility curve to get a higher stress Failure point at 70% greater than their standard design limit which is same as US.

I only use Yield in a discussion as they are findable more easily.

Thanks dd, I have no experience or expertise in this field but in the oil industry maximum working pressure the test pressure is calculated as a function of the yield however some engineering standards have it at 90% others have it at 85%. I was pointing out possibility no 1 from your post.

 

[drgondog]

I still have my LTV and P&W pocket size handbooks with a phenomenal amount of useful data, tables, plots materials/oxidizer/propellant properties.

Perusing through the example of 24S-T3 at STP has Tensile Strength (allowable/design) of 68KSI. At 400 degrees it reduces to 28KSI. For this aluminum, the allowable design stress in tension would be 68,000 pounds per square inch and Ultimate would be 1.5*68000 = 102KSI.

For high reversible load life (such as rotor pylon related structure on Huey) the allowable factor reduces 75% due to aluminum issues with fully reversible/cyclical loading. Back in my day, "Stiff'" was a good thing but high speed vibration issues have made the study of Aero elasticity much more important. For example, the A-10A fleet is getting replacement wings at DM

 

[Zipper730]

True - and the placard references Limit dive speeds as a function of IAS/Altitude so the pilot doesn't have to do the conversion.

Very true, but when you consider a given mach number corresponds with a given temperature to a given TAS, and IAS corresponds to a given TAS/Mach at different altitudes.

The structural discussion is better reduced to Dynamic Pressure encountered.

I suppose, but at the penalty of sounding slightly dense, wouldn't you use indicated airspeed and mach-number to do that?

OTOH M=1 at 30,000 feet is 589 kts and 678 mph.

That sounds right on the money

The easy calc is multiply 29.04*sqrt (T) with T in Rankin.

And that gives you TAS in knots?

Interesting - Source for German standards? Aluminum has a Failure point at approximately 1.5Yield. What made German aluminum so robust past Yield?

The source would be Captain Eric "Winkle" Brown's book "Wings of the Weird and Wonderful".

I didn't know the safety factors were based on the aluminum failure point, but merely some arbitrary safety factor because of the fact that the US used 1.5, the Germans were using 1.7, and I think the British used less than the US (interestingly, the US took advantage of this when building the P-51H)

Didnt the propellor fall off one of the spitfires used in the high speed diving test?

In one case it did, the pilot managed to glide the airplane in.

But the high mach number (0.9M) had been achieved in an earlier test without damage to the aircraft.

I was looking through World War II Aircraft Performance and the fastest dive in those tests was 0.891M.

There were later dives done to 0.92 during WWII that involved a modified PR variant with a stabilator, and eventually in 1951, a dive to 0.94. These dives were certainly in a realm that was unsafe to fly in on a routine basis.

Its a safety factor Bill. So for the design load the Germans were calculating stresses using 13% greater than in teh US, meaning that the structure has to be stronger/heavier.

They built them tougher, and low-balled the rated load?

 

[drgondog]

I suppose, but at the penalty of sounding slightly dense, wouldn't you use indicated airspeed and mach-number to do that?

Supposed to, pilots should have to calculate Q in a dive - better use the placard dive speed IAS as a function of altitude - hence the reason for flight tests to push past Limit stresses up to 'just short' of Ultimate

And that gives you TAS in knots? Yes. M1 in mph = 33.42*sqrt (T)

The source would be Captain Eric "Winkle" Brown's book "Wings of the Weird and Wonderful".

I suspect that was either a typo or he was confused for all the reasons Pbehn and I were discussing

I didn't know the safety factors were based on the aluminum failure point, but merely some arbitrary safety factor because of the fact that the US used 1.5, the Germans were using 1.7, and I think the British used less than the US (interestingly, the US took advantage of this when building the P-51H)

I suspect you are confused navigating between Design/Ultimate Stress as a function of Gross weight and the 'table stress' limit tensile load for a specific material.

There are facts the support a.) P-51 was designed at 8/12G for 8000 pounds, that b.) the P-51D never was designed to 8/12G because it would have taken a complete redesign of critical structures to beef it up to that threshold at GW=10,200 pounds and much of the structure in the D was the same structurally as the P-51, and c.) that the P-51H was in fact more robust at design 7.3/11G for 9600 pounds (more aligned to RAF standards - but at the design combat weight.



I was looking through World War II Aircraft Performance and the fastest dive in those tests was 0.891M.

There were later dives done to 0.92 during WWII that involved a modified PR variant with a stabilator, and eventually in 1951, a dive to 0.94. These dives were certainly in a realm that was unsafe to fly in on a routine basis.

They built them tougher, and low-balled the rated load?

 

[Zipper730]

Supposed to, pilots should have to calculate Q in a dive - better use the placard dive speed IAS as a function of altitude - hence the reason for flight tests to push past Limit stresses up to 'just short' of Ultimate

Because most test pilots don't feel like dying. I can't hold that against them.

I suspect that was either a typo or he was confused for all the reasons Pbehn and I were discussing

Wuzak's explanation might have been right

I suspect you are confused navigating between Design/Ultimate Stress as a function of Gross weight and the 'table stress' limit tensile load for a specific material.

There are facts the support a.) P-51 was designed at 8/12G for 8000 pounds, that b.) the P-51D never was designed to 8/12G because it would have taken a complete redesign of critical structures to beef it up to that threshold at GW=10,200 pounds and much of the structure in the D was the same structurally as the P-51

8000 wasn't the full-load, so I guess that was based on a certain fuel percentage? Regardless, the P-51D's maxes would be 6.275g/9.41g?

the P-51H was in fact more robust at design 7.3/11G for 9600 pounds (more aligned to RAF standards - but at the design combat weight.

That's an overload tolerance of 1.51

 

[Zipper730]

1. Is there anybody who has the pilot manual of any of the planes I mentioned?
2. Is there anybody who knows the maximum mach number of the following planes even if nobody has the maximum airspeed?

 

[drgondog]

1. Is there anybody who has the pilot manual of any of the planes I mentioned?
2. Is there anybody who knows the maximum mach number of the following planes even if nobody has the maximum airspeed?

I have all the P-51 manuals, the NAA Performance Analysis Reports for the P-51D-5 and P-51H, the Technical Orders, the NAA Drawing package, The 350 page P-51D Wing Structural Analysis Report, most of the published flight tests.

Maximum Mach Number is what ever that airplane achieves without losing something important. For the P-51D as tested by RAF the Mach number achieved in the dive tests was .85M and salvaged after the test due to severe damage to the wing skin (buckling), flaps, etc.

The Placard dive speed was .75M

 

[Zipper730]

I have all the P-51 manuals, the NAA Performance Analysis Reports for the P-51D-5 and P-51H, the Technical Orders, the NAA Drawing package, The 350 page P-51D Wing Structural Analysis Report, most of the published flight tests.

Maximum Mach Number is what ever that airplane achieves without losing something important. For the P-51D as tested by RAF the Mach number achieved in the dive tests was .85M and salvaged after the test due to severe damage to the wing skin (buckling), flaps, etc.

The Placard dive speed was .75M

For some reason, I thought the maximum safe was 0.80-0.84: You learn something new everyday!

 

[drgondog]

For some reason, I thought the maximum safe was 0.80-0.84: You learn something new everyday!

The placard speed is the maximum 'safe speed'.

0.80 to 0.84M is NOT a safe speed...