Bf-109 vs Spitfire vs Fw-190 vs P-51

[drgondog]

On the recognition issue, I agree that at (or near) a profile view iw would be very difficult to discern a 190D, from a Ta-152 H in combat.

When even partial plan view is seen the difference is obvious.

However I also think mistaking a Thunderbolt for a 190A is a pretty big one, comparable to the mistaking a Spitfire for a P-51 comparison soren mentioned. They are torally different in wing and fusalage shape (and size), and the only significant similarity would be the radial engine. (but even then the 190 has the huge spinner as an obvious difference)
IMO it would be easier to mistake a P-51 for a 190A. (in plan or profile the Corsar would be pretty similer to the 190 too)


And what about the Ta 152 C?

One wonders but I guess you have to be there. An awful lot of friencly fire incidents of 47s bouncing 51's and B-17 gunners shooting at both. A good friend of mine shot down what was probably a 190D, but he thought it was a 109 because it was 'an inline engine'. I think I figured it out based on his descreiption of head on pass with wing root guns firing and six o'clock view 'based' on clear 'bubble like' canopy coming away.

He (Bill Lyons) even described the scene of the head on pass on the Battleground series, Mustang.. there were quite a few 355th pilots on that one including Fortier, Lyons, Miller and Garlich.

The latter could have been an erla hood, but no 109 past the E had wing root guns.

 

[Soren]

People get back on topic please.

 

[renrich]

Soren, the Spitfire, Seafire was never robust enough to be a satisfactory carrier fight. Don't just assume the TA had a strong enough landing gear to withstand carrier landings. Perhaps you might consider sticking to landplanes and the ETO.

 

[drgondog]

Bill,

The Bf-109's Clmax figure is for the entire wing, the slats being responsible for about a 12.5% increase in Clmax, the original Clmax without the slats being around 1.51 - 1.55. This is taking into account that V24, a Bf-109F with no slats and a shortened wing span and lower wing area, was proven to have a Clmax of 1.48 in windtunnel tests at Charlais Meudon.

If the wind tunnel tests show 1.7 CLmax for the pre-stall, slats deployed, wing for full scale model - then they would be the ones to use... that's why I believe the 1.35 is more representative than the figure of 1.46 (?) that you were using earlier for the 51.

As for drag, all we have on the Bf-109 Spitfire from acual tests are the Cd0 figures:

Bf-109F G: 0.023
Spitfire: 0.0229
Bf-109K: ~0.021

Soren, do you have any one reference, preferably the Spit, that I could look at?

Nice work on the L/D but we don't need them to get to Thrust, or to look at equilibrium velocity and bank angle for 'theoretical' stall point.

 

[Soren]

Renrich,

If the Spitfire could be converted so could the Ta-152, and better so than any of the other landbased fighters mentioned here, which is my point.

The F4U was designed as carrier based fighter, it was designed keeping in mind that the rear fuselage had to endure high stresses on landing. On top of that being a carrier fight it had to have a short take off landing roll, and so it was designed with massive flaps.

Now that having been said, I will say what I've said many times by now: The F4U is undoubtedly among the top 5 fighter a/c of WW2 IMO. It was an excellent design considered what it had to accomplish and the performance it achieved on top of that. The only fighter a/c to actually top it in any significant way during WW2 was the Ta-152H, Me-262, He-162 Fw-190 D-13.

Anyway enough of this, lets get back on topic please.

 

[drgondog]

For your a+b questions: see that FW document:

Bada - thanks for the charts and references

C: don't know if that even existed. I read somewhere,( but have to find it again, i think it was in the book: Kurt Tank airplane manufacturer and testpilot) that the 190A was build to be able to sustain a max continuous load of 14G on the wings and something like20 G on the fuselage, so far above any pilot physical capbility. It seems a little bit extreme to me , anyway, let say is true, the 152 should ahave some similarities with this numbers, even if the continuous load would be slightly lower. So i'll tend to say that there was no max load. But maybe you wanted to say the load, as weight of the airplane, in this case, i would say that the H-1 was very close to it's maximal load in a take-off configuration.

That kind of design capability would be for an aircraft designed to crash as its primary mission..and way too heavy for practical application.

The 'norm' of the day for high performance Allied fighters was 8G Limit load and a factor of 1.5xLimit for Ultimate at a very specific aircraft weight and loading condition. Most of the designs would have considered the maximum pullout at SL or a rolling turn in Dive first. (The extreme dive consideration was the terminal dive in which the drag at that speed prevented further acceleration - but even that was extreme and not much was known about compressibility when these a/c were designed


i hope i could help.

You did - I think Soren and maybe Erich have also posted these but I couldn't find them.

Thank you.

 

[Soren]

If the wind tunnel tests show 1.7 CLmax for the pre-stall, slats deployed, wing for full scale model - then they would be the ones to use...

Roger, it was established on an actual a/c (109G) without prop, with flaps gear up ofcourse.

Nice work on the L/D but we don't need them to get to Thrust, or to look at equilibrium velocity and bank angle for 'theoretical' stall point.

Rgr, just wanted to illustrate the difference in L/D ratio between the a/c as it is important in sustained maneuvering and to the gliding performance of the a/c.

As to thrust, well here is the thrust figures for the Fw-190 A-9, D-9 D-12 at full boost in Kilograms (Added the Ta-152H as-well):

A-8: 1,836 kg
D-9: 2,227 kg
D-12: 2,273 kg
Ta152: 2,273 kg

 

[KrazyKraut]

Note: in this speedchart, you'll see a take off weigt of 4750 kg for the h-1, this weight is much lower than it should be, but there is a comment on this but i don't understand what it says,so if a german forum member could translate this.

It says:

"Gewicht mit halber Kraftstoffmenge!"

--> "Weight with half fuel"

And this refers to both Ta-152s in the test.

 

[Soren]

You did - I think Soren and maybe Erich have also posted these but I couldn't find them.

Thank you.

Bill,

Out of memory I think the load limit for the Ta-152H was 8.5 G at 4,750 kg and 7.7 G at 5,220 kg.

 

[drgondog]

Bill,

Out of memory I think the load limit for the Ta-152H was 8.5 G at 4,750 kg and 7.7 G at 5,220 kg.

Those figures would make sense to me as Limit Loads posted for pilots at those weights. The 51 was 8 at 8,000 pounds and I believe the load case for that was 8G dive pullout - symmetrical..

The above thrust loads are interesting but, for whatever case you want to look at, they will have to be compared against the altitude, weight and speed to attempt the manuever calculation Thrust profiles... I would speculate that they might be for sea level and were done on an engine test stand. If so, then at same RPM, prop, boost, etc the figures should vary with density.

On the CDwet from Lednicer, they are at .5 Mach as I recall.. and unfortunately are not linear with either speed or altitude.. so when the various altitudes are compared, there is this wild card to consider..I need to scratch my head and ponder whether they would be linear with respect to RN..

And as far as lednicer's sources, I don't know what the context is - relative to Reynolds Number Equivalency. I also suspect - but don't know - that these are wind tunnel tests results at or around SL

Do you recall me making a comment some time ago that these Manuever calcs are "complicated" - just for one a/c?

 

[Soren]

Good luck with it Bill !

 

[Soren]

No seriously, let me know if you need any documents, I'm stacked.

 

[drgondog]

No seriously, let me know if you need any documents, I'm stacked.

I actually need to ponder where we are 'relatively' speaking. I do not have the time or the energy to screw around trying to develop Parasite Drag for these ships.

We don't need any Parasite(Wetted, whatever) Drag test data or 'spot' data from sources such as Lednicer's Tables -

because if we have a reliable thrust figure for each of those engines at each altitude we care about - along with the velocities from the Test Report at those Hp/boost settings - at each altitude, we have all we need to calculate Induced Drag and solve for Parasite drag..

Remember that in the case of a 51B against say, a 190A-8, entering a turning fight at 25,000 feet the 51B isn't quite all out (re Thrust) entering the turn when we match these two - and ditto when a P-51D enters a turn with a Ta 152H-2, the Ta 152 will have a little thrust available as they enter at the 51D's max speed at that altitude... so the free body diagram for the Ta 152 in the latter example doesn't include the fact that the Ta 152 still has a slight acceleration capability while the 51D is all out in these equations.

In a fight this would be very important, but for purposes of the turn calcs, increasing the bank angle is more important.

 

[kool kitty89]

What speeds are you talking about for the thrust values?

 

[drgondog]

What speeds are you talking about for the thrust values?

KK -The max speeds from different flight tests at the altitudes we want to look at,, say offhand "442 mph at 25,000 feet" for whatever loading condition that particular P-51B was cited for in a specific documented test. However, we want to get thrust at altitudes in which speeds are not contained in the Report tables

As you have noted while we are scratching our heads, each of the USAAF Mustang tests conducted by the Army during WWII contained on Mike William's site have the T.O. Weight, the altitudes cites for Max Speed at peak High and Low Blower, those specific altitude 'points', the Hp based on Mfr Performance Chart, the Boost and RPM as noted by pilots for the flight profile flown.

The charts are much better for us than summary tables because they plot the continuous values as a function of altitude from SL to usually around 30,000+ feet... but Hp is not thrust, it's power.

And that power has to be translated to force to enable the free body diagram equations to be solved. A bench test value for the BMW801 and Jumo 213 that Soren just showed (at least I suspect it is a mfr 'bench test') presumably has the right, prop, peak rpm - and probably at SL... theoretically that same setting should vary by altitude with the non-linear change in density.. but I just don't recall - we didn't do much 'prop' and all my early industry stuff was with jet engines, later stuff was rotor systems and that is another world.

 

[kool kitty89]

Thrust derived from a propeller (with engine power independant of speed) has an inverse relationship with speed, correct.

(prop efficiency aside) idealy speaking:

power(kW)= thrust(kN) x velocity(m/s)

thrust(kN)= power(kW)/velocity(m/s)


And that works for many circumstances (given the prop efficiency), but the most obvious problem would be that static thrust would mean deviding by zero.


From a NASA jet a/c history page:

To give meaning to the different operating characteristics of the two types of engines, a simple example is offered as follows: A 10 000 pound propeller-driven fighter is powered by a 1600-horsepower engine and is capable of a maximum speed at sea level of 410 miles per hour. Near the beginning of the takeoff roll, the thrust at 25 miles per hour is estimated to be about 7500 pounds. Since the power is constant and proportional to the thrust times the velocity, the thrust at 410 miles per hour is about 1168 pounds. (Propeller efficiencies of 30 and 80 percent were assumed for the low-speed and high-speed conditions, respectively.) Accordingly, the thrust-to-weight ratio for the two conditions varies from 0.75 at 25 miles per hour to 0.12 at high speed. A jet fighter with the same 10000-pound gross weight and having an engine of 2500-pounds thrust has a takeoff thrust-to-weight ratio of 0.25 - and at 410 miles per hour still retains this thrust-to-weight ratio because of the nearly constant thrust characteristic of the engine. The power usefully employed in propelling the jet aircraft varies from 167 to 2740 horsepower as the speed increases from 25 to 410 miles per hour. These results are summarized in the following tabulation:

See link for data table


The results in the tabulation indicate the following two conclusions:


The thrust-to-weight ratio T/W of the jet aircraft is small compared with that of its propeller-driven counterpart at low speeds. Thus, the acceleration of the jet aircraft on takeoff will be low; and the takeoff distance, correspondingly long.
The maintenance of a nearly constant thrust-to-weight ratio through the speed range, however, gives the jet aircraft an important advantage at the high-speed end of the flight spectrum. Assuming that both hypothetical fighters considered have approximately the same drag area, the jet-powered machine would be expected to be much faster than the 410 miles per hour given for the propeller-driven aircraft. (Actually, level flight speeds as much as 100 miles per hour faster than those of contemporary propeller-driven fighters could be achieved by several of the early jet fighters.)

This ignores the problems with the zero value in static conditions.

 

[Soren]

Bill,

The thrust figures were established from bench tests conducted at SL, very correct.

KK,

You probably know the below but I'll explain it so everyone on the forum has a chance to understand.

A propeller works by accelerating air backwards faster than freestream, creating thrust. However as speed increases the efficiency of the propeller decreases, the air around the a/c (freestream) gradually catching up with the speed of the air being pushed backwards by the prop = no more thrust. Now sometime before this the prop won't be able to produce thrust enough overcome the increased drag of the a/c associated with any increase in speed = the a/c stops accelerating. So although that at the a/c's top speed the prop is still generating actual thrust, it isn't enough to offset the drag being generated.

Now for jets only the latter counts as jet engines don't loose thrust with increases in speed, and thus top speed is higher. However just like a piston engined a/c a jet a/c generates drag whilst moving through the air, and more the faster it goes, and at some point this drag overcomes the amount of thrust generated by the engines = the a/c stops accelerating. But because the jet engine doesn't loose efficiency (thrust) with increases in speed, jet a/c can go faster before drag overcomes thrust, while this happens allot more rapidly with a prop driven a/c as thrust decreases at the same time as speed and therefore drag increases.

Hope this helped some understand

 

[drgondog]

Actually KK for the problem we are working on we don't assume any propeller efficiency or deal with equations that require precise propeller efficiencies to convert Hp and velocity to Thrust.

At top speed the airplane is in equilibrium to the forces acting on it. It ain't accelerating and it's not decelerating, the engine is operating at top power, the prop is operating at whatever efficeincy its operating at - and I don't care because I need forces - not power for my equations.

Thrust = Combined Drag of Induced Drag and Parasite Drag
Push = Pull


Push is Thrust - thehorizontal Force acting on the body to put it, and keep it, in motion
Pull (Drag) is a function of ALL of the various horizontal forces on the body being 'pulled' (or pushed) through the medium (in this case - air)

Drag = Induced Drag and Parasite Drag - forces in opposite direction to Thrust

Induced Drag is a function of Aspect Ratio, the 'efficiency' of the wing plan form design, the Lift Coefficient of the wing at a specific angle of attack, the density of the air, and the Velocity (Kts).

Parasite Drag is everything else trying to hold back the airframe and keep it from accelerating. It is the drag of individual sub components like the tail, engine nacelles, interference between wing and fuselage, friction on the surfaces (depending on roughness and the velocity and whether it is in turbulent or laminar flow), gaps in control surfaces, etc, etc.

This is the dreaded 'Parasite or Wetted Drag that we have been debating of late - but I don't have the data to calculate this beast, so this is the one I want to solve for - at each altitude and equilibrium point that Soren and I have been kicking around.

So, long way around.

I can alsolutely calculate Induced Drag with small error because Soren has (correctly) retrieved the respective Aspect Ratios and the "e" for each wing, and rho (density) is obtainable right from the charts or I could calculate given STP at Sea Level for every altitude value we want to compare.

I feel we can get to precise Thrust at each altitude if we have SL bench Test data for top power with the correct prop - at every altitude we want to compare and we don't have to futz around guessing about prop or gearing efficiencies.

We can get nowhere NEAR the Parasite Drag by taking each airframe and digging around the old musty design files that may or may not exist that have All the Parasite drag data for That airplane at THose Speeds at That Altitude.

So,

Thrust

You have any data to contribute?

So, at top speed at a specified altitude I know the Knots, I know the density of air at that point and I can calculate the Induced Drag. I have two forces that I DON'T know yet, but I only need the simpler to get what I can solve for.

 

[drgondog]

Bill,

The thrust figures were established from bench tests conducted at SL, very correct.

KK,

You probably know the below but I'll explain it so everyone on the forum has a chance to understand.

A propeller works by accelerating air backwards faster than freestream, creating thrust. However as speed increases the efficiency of the propeller decreases, the air around the a/c (freestream) gradually catching up with the speed of the air being pushed backwards by the prop = no more thrust. Now sometime before this the prop won't be able to produce thrust enough overcome the increased drag of the a/c associated with any increase in speed = the a/c stops accelerating. So although that at the a/c's top speed the prop is still generating actual thrust, it isn't enough to offset the drag being generated.

Now for jets only the latter counts as jet engines don't loose thrust with increases in speed, and thus top speed is higher. However just like a piston engined a/c a jet a/c generates drag whilst moving through the air, and more the faster it goes, and at some point this drag overcomes the amount of thrust generated by the engines = the a/c stops accelerating. But because the jet engine doesn't loose efficiency (thrust) with increases in speed, jet a/c can go faster before drag overcomes thrust, while this happens allot more rapidly with a prop driven a/c as thrust decreases at the same time as speed and therefore drag increases.

Hope this helped some understand

Great summary - propeller and rotor/recip engine design is another 'arcane' (did not say witchcraft-lol) field of aero engineering especially when they were actually trying to get close to 550TAS with prop/aircraft systems

 

[buzzard]

Soren, drgondog, and kk89,

This is all very informative and interesting. Great stuff from all of you

Thanks,

JL