Stiffening of Controls for the A6M/A7M series

[Irregualr23]

Direct comparison of elevator angle deflection of Zero with other planes doesn't make sense since same American reports you use say that the records from the elevator control system's angle sensors are not the actual elevator deflection angle because Zero's control system stretches to change the "transfer torque" from the stick to the elevator at different speeds by stretching the control cables.
The discrepancy in the sensor reading was caused by the fact that it was located near the cockpit and not on the elevator itself. Stretching of the control system cable leads to the actual excess of sensor readings relative to the actual deviation

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It is indeed a point to check out, 18deg delfection to reach CLmax seems to be excessive. However the control cable on Zero, as quoted by someone else above, was made to be stiff so that no much room to get stretched, which was also an important factor for its large stick force at high speed. There was an entire essay in Japanese explaning how the absence of elasticity in Zero's control cable lead to high control force.

Maybe I don't understand what you mean but if plane produce positive lift power on tail horizontal with classic aerodynamic scheme - it means plane made with unstable scheme where center of mass located behind the aerodynamic pressure center. This thing is common for modern jet fighters like Su-27 but not the WW2 props which was made with stable scheme

And because of stable scheme this issue:

WW2 was a point where we see some fighter planes to have loose stability margin, and especially the propeller driven by high-output engine in the front couples with aircraft's longitunal stability, usually distabilize the aircraft. The fact that most late-war Allied fighters had their C.G. placed after the aerodynamic center of wing-fuselage combination, but forward of the neutral point (i.e. the aerodyanmic center for wing-fuselage + empennage), which means they are still static-stable design. In most cases, these airplanes would only require ~5 deg of up elevator deflection to reach CLmax, and at this AoA, most of time the h-tail would generate slightly positive tail lift.

The postive tail lift doesn't necessary require an unstable C.G. location, but with mordern fighter having fly-by-wire they can have very small stability margin that the tail plane would deflect down in the manuever to generate significant positive tail lift.

 

[skultew]

However the control cable on Zero, as quoted by someone else above, was made to be stiff so that no much room to get stretched, which was also an important factor for its large stick force at high speed. There was an entire essay in Japanese explaning how the absence of elasticity in Zero's control cable lead to high control force.

According to Jiro Horikoshi book and other books Zero control cable was specially made stretchable rather than completely rigid so I honestly don't really understand where the claims that the cable was rigid come from

WW2 was a point where we see some fighter planes to have loose stability margin, and especially the propeller driven by high-output engine in the front couples with aircraft's longitunal stability, usually distabilize the aircraft. The fact that most late-war Allied fighters had their C.G. placed after the aerodynamic center of wing-fuselage combination, but forward of the neutral point (i.e. the aerodyanmic center for wing-fuselage + empennage), which means they are still static-stable design. In most cases, these airplanes would only require ~5 deg of up elevator deflection to reach CLmax, and at this AoA, most of time the h-tail would generate slightly positive tail lift.

It is pretty interesting, do you have some good reports where more detailed information can be read?

But some doubts appear because if US planes had center of gravity and aerodynamic center so close this means that planes should be pretty unstable and their controls should resemble the I-16 with all the known problems, doesn't it

 

[Airborne2001]

From what I've read, I think the giant non-boosted ailerons on the A6M didn't help with the compressibility issue. The I-16 had this problem too. These giant ailerons were great for low speed dogfighting, but at high speeds they were a huge detriment, especially when they weren't boosted.

EDIT: Plus, the A6M (especially prior to the M7) being made of such thin skinning didn't help.

 

[Irregualr23]

According to Jiro Horikoshi book and other books Zero control cable was specially made stretchable rather than completely rigid so I honestly don't really understand where the claims that the cable was rigid come from

I see what's happening, I remembered it in the opposite way, because IJN had a requirement for rigid cable while Horikoshi did the opposite. This ensures the elevator to be less sensitive especially at high speed. Which also explains the high-stick force persisted as late version of A6M5c moved C.G. to >27% MAC.

Similar system was also applied on N1K2 in which the pilot would select the gear ratio of the elevator cable for high-low speed mode. As the NACA instrument was not adjusted for this elasticity, the recorded elevator deflection anlge was exaggrated, also in one of the report, this large deflection angle was quoted to show "excellent longtinual stability".

It is pretty interesting, do you have some good reports where more detailed information can be read?

A typical aircraft C.G. vs. A.C. looks like below, C.G. been placed after the A.C. of wing-fuslage combination:

The longtitunal stability margin was provided by the tail volume, which in simple can be understood as Area of h-tail x H-tail Arm distance. In particular, F6F and A6M were two aircrafts that had very large tail-volume among WW2 fighters. This had provided extra longtinual stability, even with C.G. been placed rearward. F6F had Area-tail/Area-wing ratio of 23%, A6M had around 22%, Ki-84 ~ 19.8%, F4U ~ 18.4%, FW-190 ~ 15.6%.

I have a report for small-scale XF6F-1 wind-tunnel on revising H-tail:

With C.G. ~26.42%MAC, with no H-tail the dM/da is positive (unstable), which means C.G. 26.42% is after the A.C. of wing-fuselage. With tail been placed, incidence angle and tail area enlarged, dM/da becomes negative which means the aircraft is stable.

We can calculate with F6F's CL linear coefficient of 0.07 at this reynold number scale, the calculated A.C. for F6F's wing and fuselage is located around 22%MAC. With tail been placed, the aerodynamic center moves rearward due to tail lift to around 35-37% MAC, provides 8-13% MAC stability margin for the aircraft, under power-off gliding condition.

My understanding is that, a typical WW2-type aircraft with elevator de = -5 deg and would have wing CLmax ~ 1.25, fuselage contribution ~0.15, and tail lift 0.05, makes total trimmed CLmax 1.45. If de = - 15deg is required, for example the aircraft is overly stable, then tail contribution becomes -0.05, aircraft CLmax dropps to 1.35. If the aircraft is unstable and uses positive de = 5 deg in turning, tail lift increases to 0.15 this boosts trimmed CLmax to over 1.55. This demonstrates how stability-margin and tail lfit affect the aircraft's manueverability, though deflection angle is also affected by the Cm of the wing and tail volume. In the previous case, the relatively small deflection angle for F6F can be achived by larger tail-volume and using low-pitching-moment airfoil such as NACA23012.

But generally, you wouldn't expect an WW2-era fighter to squeeze too much lift contribution from the tail, most of time the effort was made to avoid the loss of lift due to negative h-tail lift with elevator deflected up. (A6M Zero's narrow elevator may used to focus more on elevator-hinge moment rather than trimming using reduced H-tail lift, which is also a thoguht to avoid lossing too much tail lift)

Of course, if the aircraft becomse too unstable, the extra boosted CLmax may be unusable since the aircraft would be hard to control in the accelerated flight. Which is why reason both IJN and NACA test pilots actually preferred Zero's flight qualities on large longitunal stability.

But some doubts appear because if US planes had center of gravity and aerodynamic center so close this means that planes should be pretty unstable and their controls should resemble the I-16 with all the known problems, doesn't it

You are right, lots of Allied single-engine aircrafts had to face longitunal stability problem with overloaded C.G. position and power-on condition. Famous example such as Spitfire, and P-51D with rear fuselage tank filled, and surprisingly, F8F, whose longtinual stability becomes neutral when power is added above the trimmed airspeed. I would say probably the Allied fighter pilots got used to aircraft with less stability-margin compared to Japanese and Russian pilots. Of course, the case of I-16 was caused by even more afterward C.G. of 33%-35% MAC, with low-moment airfoil Clark Y-H been used, the problem becomes more significant.

Russian aircratfs also tended to place C.G. forward, usually in front of 24% MAC for both Yaks and LaGG-3/5/7. On the LaGG-3 the Russian engineer also placed a weight balance intended to make control heavier to prevent over-stressing aircraft at high-speed --- the very similar reason Horikoshi reduced the control elasticity of the Zero.

 

[skultew]

View attachment 883099
I see what's happening, I remembered it in the opposite way, because IJN had a requirement for rigid cable while Horikoshi did the opposite. This ensures the elevator to be less sensitive especially at high speed.

Exactly.
And because of it, as I said and as written in US reports - data about elevator deflection angle during US tests can't be used because it is incorrect. It was the main point this conversation started.

Which also explains the high-stick force persisted as late version of A6M5c moved C.G. to >27% MAC.

A6M5 Hei center of gravity with full fuel and ammo was 28,4% MAC and 29,8% MAC with 250kg bomb

Similar system was also applied on N1K2 in which the pilot would select the gear ratio of the elevator cable for high-low speed mode.

Yes, It is pretty interesting system. I didn't know should there be written about it but since we've already touched on this topic

As it was said the elastic control system did not meet IJN requirements and these requirements were not amended when the Zero was accepted.
Instead it was decided that Zero would receive special permission to operate with this feature.
It was assumed that the company would have to apply for permission to use elastic cables but no other company besides the Mitsubishi with Zero applied for this.
Instead of this system Kawasaki made system with mechanical changed control forse

As the NACA instrument was not adjusted for this elasticity, the recorded elevator deflection anlge was exaggrated, also in one of the report, this large deflection angle was quoted to show "excellent longtinual stability".

From a certain perspective, this could indeed be called stability.
The US report considers the aircraft to be purely an external contour with corresponding profiles and centers of mass, but this control system design did give the pilot a greater sense of control over the aircraft, which is precisely what it was intended for.
In one book suggested that the Americans never fully understood the meaning of the elastic control cable although they certainly noticed the fact of this elasticity.

A typical aircraft C.G. vs. A.C. looks like below, C.G. been placed after the A.C. of wing-fuslage combination:

The longtitunal stability margin was provided by the tail volume, which in simple can be understood as Area of h-tail x H-tail Arm distance. In particular, F6F and A6M were two aircrafts that had very large tail-volume among WW2 fighters. This had provided extra longtinual stability, even with C.G. been placed rearward. F6F had Area-tail/Area-wing ratio of 23%, A6M had around 22%, Ki-84 ~ 19.8%, F4U ~ 18.4%, FW-190 ~ 15.6%.

I have a report for small-scale XF6F-1 wind-tunnel on revising H-tail:

With C.G. ~26.42%MAC, with no H-tail the dM/da is positive (unstable), which means C.G. 26.42% is after the A.C. of wing-fuselage. With tail been placed, incidence angle and tail area enlarged, dM/da becomes negative which means the aircraft is stable.

We can calculate with F6F's CL linear coefficient of 0.07 at this reynold number scale, the calculated A.C. for F6F's wing and fuselage is located around 22%MAC. With tail been placed, the aerodynamic center moves rearward due to tail lift to around 35-37% MAC, provides 8-13% MAC stability margin for the aircraft, under power-off gliding condition.

My understanding is that, a typical WW2-type aircraft with elevator de = -5 deg and would have wing CLmax ~ 1.25, fuselage contribution ~0.15, and tail lift 0.05, makes total trimmed CLmax 1.45. If de = - 15deg is required, for example the aircraft is overly stable, then tail contribution becomes -0.05, aircraft CLmax dropps to 1.35. If the aircraft is unstable and uses positive de = 5 deg in turning, tail lift increases to 0.15 this boosts trimmed CLmax to over 1.55. This demonstrates how stability-margin and tail lfit affect the aircraft's manueverability, though deflection angle is also affected by the Cm of the wing and tail volume. In the previous case, the relatively small deflection angle for F6F can be achived by larger tail-volume and using low-pitching-moment airfoil such as NACA23012.

But generally, you wouldn't expect an WW2-era fighter to squeeze too much lift contribution from the tail, most of time the effort was made to avoid the loss of lift due to negative h-tail lift with elevator deflected up. (A6M Zero's narrow elevator may used to focus more on elevator-hinge moment rather than trimming using reduced H-tail lift, which is also a thoguht to avoid lossing too much tail lift)

Of course, if the aircraft becomse too unstable, the extra boosted CLmax may be unusable since the aircraft would be hard to control in the accelerated flight. Which is why reason both IJN and NACA test pilots actually preferred Zero's flight qualities on large longitunal stability.

But some doubts appear because if US planes had center of gravity and aerodynamic center so close this means that planes should be pretty unstable and their controls should resemble the I-16 with all the known problems, doesn't it

You are right, lots of Allied single-engine aircrafts had to face longitunal stability problem with overloaded C.G. position and power-on condition. Famous example such as Spitfire, and P-51D with rear fuselage tank filled, and surprisingly, F8F, whose longtinual stability becomes neutral when power is added above the trimmed airspeed. I would say probably the Allied fighter pilots got used to aircraft with less stability-margin compared to Japanese and Russian pilots. Of course, the case of I-16 was caused by even more afterward C.G. of 33%-35% MAC, with low-moment airfoil Clark Y-H been used, the problem becomes more significant.

Russian aircratfs also tended to place C.G. forward, usually in front of 24% MAC for both Yaks and LaGG-3/5/7. On the LaGG-3 the Russian engineer also placed a weight balance intended to make control heavier to prevent over-stressing aircraft at high-speed --- the very similar reason Horikoshi reduced the control elasticity of the Zero.

I am glad for all this description of the topic. But I understand that need to research some stuff to give the good and thoughtful answer. Hope will do it soon