coriolis effect & static droop [Archive]

212man

12th Jan 2001, 03:31

I think you'll find that what Offshoreigor is referring to is actually transient droop, so named because it is not permanent. As he says, the governing system senses N2 variation (or Nr if piston) and makes and adjustment to restore it to a preset value. To assist with this most systems use some input from the collective, either mechanical (droop compensation cam) or electronic (anticipator) to'forewarn' the system of an impending change and to start adjusting before the change occurs, so minimising the effect.

Static droop is when the Nr decays to a new value after application of power, and does not return to the governed speed. This may be due to a fault if pronounced, however some degree of in-built droop is a requirement of hydromechanical systems to aid their stability. ie to prevent them oscillating about the datum. I'd like to say more, but it's late, I'm tired and I'd prefer to quote from more authorative sources to ensure accuracy. If you can get access to Simon Newman's book on Helicopter Aerodynamics, he has a good chapter al about Turbine engine governing.

Regarding the coriolis effect, don't forget that it is more concerned with the flapping of the blade in forward flight ie the advancing blade rises, so speeds up then slows down as it flaps down on the retreating side. Hookes joint effect has a similar effect as do periodic drag changes.

Hope this tallies with your understanding.

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Another day in paradise

Lu Zuckerman

12th Jan 2001, 05:09

To: Helimutt

Corriolis force (conservation of angular momentum) is the force that causes the blades to lead and lag. The change in the blade mass is often diagrammed out showing one blade in various angular positions relative to the flapping hinges. The diagram shows the relationship of the blade mass at each position relative to the rotational axis. The diagram only shows one blade flapping up and the other blade is in the pure radial position and this can’t be. Another means of explaining this phenomenon is to have you picture an ice skater that brings her arms closer to her body and as a result speeds up in rotation and when she moves her arms outwards slows down. Speeding up and slowing down is due to her changing he center of mass.

Regarding what your instructor stated about the small change in blade mass CG not being sufficient to cause the leading and lagging I think he is not considering the blade mass due to centrifugal force. On a large helicopter the blades have a very large weight due to centrifugal force and could weigh in excess of 70,000 pounds at the lead lag hinge. So at this weight it does not take a very large change in mass CG to exert the energy to move the blades about the lead lag hinge.

Now, this is the way I learned it at Sikorsky over 45 years ago. It is difficult to explain without drawing diagrams but here goes.

First, something to consider: Discounting mast tilt to compensate for tail rotor drift or forward tilt on a Sikorsky mast that adds forward cyclic, visualize the disc as being horizontal and rotating about a fixed axis. This is the drive axis. In order to have lead and lag you must have an offset hinge. On the Robinson the cone hinges are the offset hinges. For arguments sake the rotor head does not move relative to the mast (difficult if you think of a Robinson head. Assume that the blades are coned up at the flapping hinge when you pull in collective. This will form a “V” with a flat bottom. Now lets look back at the explanation above that describes the blade in the flapped condition(s) and showing the blade mass CG moving closer to the center of drive creating forces that cause the blade to speed up or in other words, lead. This is totally false, as the blades do not lead when in a hover discounting tilt due to tail rotor compensation.

The blades lead and lag only when you introduce cyclic pitch change. Lets go back to the “V” with the flat bottom. When cyclic pitch is made the disc will tilt in the respective direction. This will make one leg of the flat-bottomed “V” move up and the other down in relation to the rotor head. Initially we discussed the drive axis. That was when you had a flat-bottomed “V”. Once the cyclic input was made the disc was displaced and now the blades are rotating about the driven axis. Due to the laws of conservation of angular momentum the blade wants to move forward from its’ position about the drive axis to the position scribed by the rotation about the driven axis. Since the blade can’t move linearly because the lead lag hinge anchors it, it will do what it can under the restrained condition. The tip of the blade will move forward on the advancing side. Now, when that same blade is on the longitudinal axis the forces diminish and the blade goes to the radial position and it wants to stay in that position. So, when it passes over the longitudinal axis, the blade will tend to hang back or actually move slower than it is being driven due to the change in mass CG position, which is lower. This is per the Ice skater illustration. It will hang back until it is over the left side of the helicopter and then return to the radial position. If you could look down on the tilted disc a four-blade rotor system would look something like a peace symbol.

While my wife and I went out for Chinese the other guys chimed in and covered droop as it applies to the rotor and the engine but there is another type of static droop and it applies to the drooping of the ailerons on a wing. When I worked on the CL604 and the Regional Jet flight control systems I discovered that after the flight control system is rigged and the Power Control Units (Servos) are turned on the pilot valves in both aileron servos are biased to move the ailerons down from the rigged position. I believe this was to both provide a bit more of lift and that aerodynamic forces would tend to move the ailerons to the rigged position and stiffen the system because the servos were always trying to get them to extend to the drooped position.

To:212man

Regarding your comment below:

Regarding the coriolis effect, don't forget that it is more concerned with the flapping of the blade in forward flight ie the advancing blade rises, so speeds up then slows down as it flaps down on the retreating side. Hookes joint effect has a similar effect as do periodic drag changes.
Hope this tallies with your understanding.

The advancing blade flaps down and the retreating blade flaps up. According to your descrioption you would be flying backwards. Your comparison to a hookes joint is valid. That is why they have constant velocity joints on a front drive car.
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The Cat

rotorque

12th Jan 2001, 08:49

Sorry Helimutt, I am just going to throw a wobly in on this thread, hopefully it will get back to your topic reasonably soon.

Lu: I have heard you quote the flapping arument before, the one that involves the leading blade flapping 'down'. I am under the impression that you may be refering to the result of the cyclic reaction that occurs in forward flight only. But in this scenario we must seperate the reactions taking place. I believe that 212man is correctly refering to the flapping 'up' of the advancing blade due to the increase in lift resulting from the increased forward speed when the blade advances. If you think in basic terms (which is the only way I can do it) The whole reason we have a reaction from conservation of angular momentum is because as the advancing blade flaps up the CoG moves inboard and the blade ties to lead, this leading may seen as a flexing of the blade or an actual movement from a hinge or a bearing. On the other side of the disk we have the retreating blade reducing its lift due to slower relative airspeed and therefor it flaps 'down', this in turn throws the CoG outward and the blade tries to slow down therefore we get the lagging of the blade. This is why we see it as a peace sign when viewed from above.
All of the above is refering to forward flight but Lu mentioned that we won't see the same sort of things in the hover. We do, especially when we pull in pitch. The blades cone up and we get conservation of angular momentum, seen as a leading of all the blades at the same time, the only problem is that at the same time the Nr is trying to droop through drag so we don't get any benefit from the reaction. The most important thing to remember in this scenario is that the reaction is only taking place while there is displacement (movement) of the blades from a previous position.

One last thing for Helimutt. You will and do see a reaction in flight when we get an increased or decreased load on the disk. Especially now that you guys are flying Robbinsons with govenors fitted. Take a look at the manifold pressure gauge when you get some gusty conditions, notice the change in power as the loads increase and decrease with the gusts, this is due to the governor working to maintain a constant RRPM. The coning angles change as a result of the gusts and therefore the RRPM tries to change as well (don't get me wrong there are heaps of things going on but this is highlighting what we are talking about).

Any way Helimutt, this thread should be fun from here. Hope you can keep up.

Cheers mate.

212man

12th Jan 2001, 16:17

Thanks rotortorque.

Lu, find yourself a photograph of an S61 or HH 53 or similar. Try and find one with the blades 'frozen': you will see the advancing blade flapped up with low pitch and the retreating blade flapped down with lots of pitch QED.

I also think, unless I've mis-read your description, that what you are actually describing is Hookes-joint effect, and not coriolis effect.

If you want to be pedantic, the blade mass does not change, nor does its weight, only the forces exerted on it.

Anymore from anyone on droop? If my PC skills were up to it I'd show static droop curves. These are produced on a test bed with the engine at varying speeds and then the N2 plotted vs load, to produce a curve. The steeper the gradient, the less the droop. This will result in hard damping though which can lead to over swinging of N2. A shallow curve results in slow return to the datum. It is possible to artificially alter the effect mechanically to have soft damping coupled with minimum droop.

Having said all that, the static droop to which the question refers is almost certainly that commonly found day to day, where, for instance, you set 100% Nr at flat pitch then lift in to an 80% Q hover and find the Nr is 98% and needs beeping up. On the 76 we set 101% Nr before lifting and this gives 100% in the hover.

Anyone with more detail to offer? it's a while since I read up on this.

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Another day in paradise

[This message has been edited by 212man (edited 12 January 2001).]

helimutt

12th Jan 2001, 22:48

Thank you gentlemen for enlightening me with your wisdom. It's going to take me a while to get this all straight in my head but im off out tonight to see if I can cause my own type of droop with a few shandies!!!
My question wasn't meant to cause arguments but it seems to have gotten everyones grey matter working. The actual question I was asked about static droop was as follows:-
Static droop results from:-
a/. increasing collective pitch
b/. decreasing collective pitch
c/. changing the cyclic pitch

Am I missing something here? Are we all talking about the same thing?
I sit in anticipation!
Thanks guys.

ShyTorque

13th Jan 2001, 01:03

Helimutt,

Quite simply, your answer is a).

Put simply, it refers to rotor RPM (Nr). More collective pulled = more static droop. This is very apparent on the older AS-330J Puma(with a mechanical FCU using flyweights and needle-valves to control fuel flow) where the flat pitch Nr is higher than the hover Nr.

However, on later versions such as the AS-332 Super Puma the Makila FCU is more sophisticated. At flat pitch on the ground the Nr sits slightly low and is artificially increased by the FCU when the lever is raised for lift-off to the hover.

We can now all discuss which way is positive droop and which way is negative droop

212man

13th Jan 2001, 02:03

Oh God, all that effort over a question like that?

ShyTorque is referring to the artificialy altered droop characteristics I alluded to. The inherent nature of the system is to droop along a curve, however the gradient can be altered to give more favourable characteristics, and even made positive as indeed the 332 is ( I think it goes from 260 Nr at MPOG to 265 in the hover?).

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Another day in paradise

Cadfael

13th Jan 2001, 02:48

While everyone here is in the 'govnr' mode. A quick question on B206 gov. Which gov do you consider better for droop compensation characteristics. (ie stays close to 100% throughout collective increases and decreases) I remember flying Seiko fitted 206's eons ago that were pretty good. These days (post AS332/350 driving) they seem more 'lazy'. Bendix or Seiko guys 'n gals ??

helisphere

13th Jan 2001, 05:15

Yes there is a reason for the increase in rotor rpm when increasing the load factor on the disc in maneuvering other than coriolis effect. The best way to make sense of it is: say you are in auto rotation. As you are gliding down if you could some how add weight to the aircraft the rpm would increase and you would have to start pulling pitch to keep the rpm constant as you take on weight. As you loaded a larger amount of weight you would get to the point where your collective would be pulled pretty high. Adding weight to the helicopter is basically the same as increasing the load factor by turning or pulling aft cyclic. Increasing load factor in power on flight does the same thing to rpm as it does in unpowered flight. RPM does not increase just because of coriolis effect. Coriolis effect only increases rpm for the time when the coning angle is increasing, once the coning angle stabilizes there is no more coriolis effect and the rotor still must overcome the same amount of drag as before the coning was increased if it is to maintain rpm. If coriolis was the only force trying to increase rotor rpm, then when you rolled into a turn torque would momentarily decrease and once you were stabilized in the turn torque would return to where it was. We pilots know this does not happen.

Now for the explanation. It will help if you dig out a helicopter book that shows the airfoil section of a blade with all of the lift and drag vectors. They should show one for powered flight and one for autorotation. If you look at the one for autorotation you will see that the total force vector is perfectly vertical. That means that there is nothing opposing rotation and nothing to increase it. Now look at the diagram for powered flight, you will see that the total force vector is tilted aft trying to slow the rotor down. The torque generated by the engine of course overcomes this. Now imagine what happens to the relative wind vector if you increase the load factor. It will become more upward actually increasing the angle of attack. But look what happens to the total force when you change the relative wind.

Remember the drag vector is always parallel to the relative wind and the total lift vector is always 90 degrees to it. The total force is the lift and drag added together.

So if the relative wind is more upward then the total force is more forward, whether it is powered flight or unpowered.

Bottom line: when you increase load factor at any collective setting, you change the relative wind to be more upward and that angles the total force forward which is a decrease in drag and rpm increases or torque is reduced in a governed engine.

lmlanphere

13th Jan 2001, 08:02

helisphere, if you are saying that once the coning angle stops increasing the rotor rpm drops back to its original value, I'd have to disagree. Nr should remain at its accelerated value till the coning angle decreases. perhaps I misunderstood.

[This message has been edited by lmlanphere (edited 13 January 2001).]

Lu Zuckerman

13th Jan 2001, 08:54

To:Rotorque, Helisphere and 212man

This might get a bit mixed up because intermixed with your comments about corriolis forces you were addressing static droop.

One of you made a comment about leading of the blades when you pulled collective in countering a statement I had made about a diagram. Whoever made that statement was correct. The leading would be limited and controlled by the amount of blade cone. If the blades were returned to flat pitch the lead would diminish to the point that the blades were in the normal position. It would be assumed that the blades in the flat pitch condition would be in the pure radial position and assuming a four-blade system the blades would form a cross. In actuality, due to the inertia of the blades instead of the blades forming a cross it would be more like a very mild X and as such the blades all would lag behind the pure radial position (Cross) and all leading and lagging would be behind the radial position. The only time the blades ever cross the radial position is during auto rotation when the blades are driving the system as opposed to the other way around.

Regarding the leading of the blades causing the rotor system to speed up (for a very short time) I don’t think this is true. Because the blades are mounted on a hinge and connected to the rotor head by a soft damper any movement of the blades about the lag hinge would not be reflected on the rotorhead. That is, unless the rotor system is in autorotation when the dampers are bottomed out and the aerodynamic forces on the rotorhead are back driving the gearbox.

If what was said above were true then the leading and lagging of the blades would be reflected in the dynamic system and the power train and be very noticeable by the pilot. The only time the pilot is aware of this condition is when he has a bad damper and he feels it as a lateral beat. But although the lateral beat stems from leading and lagging it really is a mass balance problem as the effected blade is out of balance with the other blades relative to their respective positions during rotation

Regarding the terms flapping up and flapping down there is a problem of interpretation. These terms are usually discussed in teaching helicopter aerodynamics and it causes many many problems. It is true on a two bladed rotor system if one blade flaps up the other will flap down. Assuming you are in a hover in a Bell and discounting any cyclic adjustments for tail rotor translation here is what happens. In a hover the blades are disposed at 90-degrees to the rotor shaft. When the pilot pushes forward the aerodynamic and precessional forces will cause the disc to dip down over the nose and rise up over the tail. In order to do this the advancing blade must flap down over the nose and due to the construction of the system the other blade flaps up over the tail. Now to what you were referring to. The advancing blade meets the oncoming relative wind and flaps up and the other blade flaps down. When the rotor system flaps relative to the fixed swashplate there is pitch coupling and this restores the blades to the commanded position therefore any tendency to flap out of the commanded disc position the blades immediately return. However when they use the flap up/flap down illustration they never address pitch coupling.

Regarding the comment about seeing the blades flap up on the advancing side and flapping down on the retreating side of an S-61, I would advise you to look at the tip path change when you move the cyclic. The disc will tip down in the direction of cyclic movement and, although you can’t see the rear of the disc I will assure you if one part goes down the opposite side of the disc will go up. Even on a 7 bladed CH-53.

In one post I mentioned Hookes joint effect in agreeing with a post made previously. On a hooks joint there is not a constant velocity in the rotation of the input Vs the output and that is when I mentioned a constant velocity joint being employed on a front drive automobile as opposed to using a Hookes joint in the drive system. The leading and lagging is a function of the law of conservation of angular momentum or as it is known as Corriolis forces. To minimize the effects of Corriolis forces on a single rotor helicopter they underselling the rotor system to lower the center of mass relative to the drive point on the rotor shaft. Speaking of center of mass, the center of mass on the blades does not move in or out relative to the root or tip of the blade. However it can move up or down relative to the flapping hinge and this is what causes the Coriollis force to kick in. The mass of the blades increases due to centrifugal force and does two things it increases the stiffness of the blade and it increases the moment of inertia of the disc system providing the forces necessary to cause precession when a perturbing force is appplied.

Now lets discuss the exceptions to the statements above. I previously stated that with a soft inplane rotor system the leading and lagging forces are not transmitted to the rotorhead and then into the drive system. A major exception to this is the Robinson head. The Robinson head has cone hinges, which are in effect, offset hinges that allow flapping in relation to the rotorhead itself. As I stated previously when a blade flaps it leads and lags. The Robinson head does not provide for leading and lagging brought about by flapping. Nevertheless, the tendency to lead and lag is there. These forces are so great that any lead lag motion is reacted by the cone hinges and then into the teeter hinge and then into the drive shaft and transmission. This fact is borne out by the high replacement frequencies for the respective hinges/bushings as they are worn into an egg shape. The Robinson head like the Bell head is underslung to minimize the tendency to lead and lag but the Robinson head unlike the bell head has the ability to flap and the Bell doesn’t.

Hopefully, I have covered all of the points in your respective posts. If not let me know as I am not going anywhere.

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The Cat

Grey Area

13th Jan 2001, 18:31

Concur with offshoreigor. To test this a droop plot can be taken, plotting Torque against RRPM. It is generally accepted to be normal for the RRPM to drop as TQ increases, but in these days of FADEC is is quite possible to have no change (or even an increase) in RRPM with TQ increase, should the designer wish.

One could argue that a drop in RRPM will increase coning angle as will an increase in TQ, which may not be totally desirable, especially on a rigid head - answers on a postcard to .........

Grey Area

13th Jan 2001, 18:41

And another thing. Coriolis and Hookes Joint effect are both ways of looking at the conservation of angular momentum.

GA

212man

13th Jan 2001, 21:51

Grey Area,
I rather thought your first post was echoing my sentiments?

I would think that any tendency for Nr to rise with increased coning from droop would be masked by the droop itself. ie the droop ing would not be as great as it otherwise would be. (that's how I explain it to my wife, anyway).

Lu, please avoid using expressions like "the blade mass will increase".

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Another day in paradise

Lu Zuckerman

13th Jan 2001, 22:42

To: 212man

I used the term mass as it applies to the definition of a Newton. In this case mass = weight. What I was trying to imply is that the reflected weight of a rotor blade that weighs 234 pounds would due to centrifugal force be equal to a weight of 72,000 pounds measured at the offset hinge. This weight combined with the other blades creates a disc that has a polar moment of inertia of 360,000 pounds, which provides three things. 1) Stiffness of the blade and the disc. 2) Gyroscopic rigidity of the disc and, 3) the inertial energy to cause the blade disc to move when perturbed by an external force (control input).

Hopefully in this description I didn’t misuse some words such as polar moment…..

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The Cat

212man

14th Jan 2001, 02:15

Well, whatever, the mass remains constant; it's the force that changes. As polar moment of inertia (I) is equal to a half mass times radius(squared) I suppose it rather depends on the diameter and number of blades of the system you quote figures for.

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Another day in paradise

ACORN

14th Jan 2001, 18:30

Only read halfway through this thread and started to get confused myself.

Just so I can clarify my understanding of the different topics, conservation of angular momentum surely does not relate to lead and lag, it applies to flare effects with regards increase of NR during the flare.
I was under the impression that lead and lag was as a result of the differing drag produced around the disc given changes in relative rotational velocity as the disc moves through the air.

As for static droop, it is caused by increasing collective pitch from MPOG to hover power where it is not compensated for in a governor. Any further changes in NR in flight caused by manouevering would be termed transient droop.

One way to explain conservation of angular momentum when related to flare effects is as follows:-

Imagine a flat rotor disc with an arbitary rotor blabe length which would produce a set figure for the circumference of the disc area. If you apply 'G' loading to the disc it will increase the coning angle of the disc i.e. in the flare. As the blade cones it slightly reduces the radius of the disc or the distance of the tip path plane from the shaft axis. This therefore gives a slight reduction in the circumference of the disc area as the radius is decreased. The blades momentum means it wants to travel the same distance round the disc in the same time but has a smaller disc round which to travel therefore it will travel further in terms of rotation whilst in the flare. The result is an increase of NR.

Feel free to educate me if I am too far from the mark, I'm always learning too!

lmlanphere

14th Jan 2001, 23:18

acorn, profile drag does account for some lead-lag, but there are other reasons for the phenomenon. Conservation of angular momentum is one of these key reasons.

212man

15th Jan 2001, 02:41

Periodic drag variations do cause lead and lag, as I mentioned earlier. However, coriolis effect caused by the flapping of the blades as they advance and retreat, so moving their c of g inwards and out relative to the axis of rotation, is one example of conservation of momentum.

Coning during a flare may cause an increase in Nr, however most is from aerodynamic considerations (induced flow and relative flow vectors changing lengths etc, too late for that now).

If the Nr in the hover is different to that at MPOG, then it is an example of static droop from a poorely set up governor, but some static droop is inherent and necessary in simple systems, though largely masked.

Other variations are only transient if they are not permanent and return to the previous value.

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Another day in paradise

Lu Zuckerman

15th Jan 2001, 05:12

To: Acorn

212man pretty much described how the blades speed up when you do a flare during autoratation. It is purely aerodynamic and is in no way related to the forces that cause lead and lag. Many years ago lead and lag was known to exist but was not very well understood. Sikorsky performed an experiment to determine what the blades did during the climbing and diving associated with the input of cyclic pitch. What they did was to put a movie camera on top of an S51 rotorhead that was mounted in a whirl stand. The camera was aimed from the top of the rotorhead and was in line with one blade, which was in a pure radial position. What they were looking for was the flexing of the blade in flight and what they found was very frightening. The best way to describe the action of the blade was to compare it to a sinusoidal wave that traveled down the blade until it was reacted at the spider (Flapping hinge). As an aside, when they showed this to helicopter pilots many of them gave the job up and went on to other things. This is still true on even the most modern helicopters. How this was countered can be the subject of another thread.

They received a bonus when they shot the film. The first thing they saw was that the blade left the pure radial position and sort of hung back behind the radial line. When they pulled collective the blade moved slightly forward but still stayed behind the radial line. The conventional wisdom at that time was that the advancing blade lagged due to the increased relative wind and that the retreating blade lead because of the assisting air loads. When they introduced cyclic the advancing blade lead and the retreating blade lagged. At no time did the blade move past the radial line even with severe cyclic input.

In response to one of your statements conservation of angular momentum has everything to do with lead and lag.

In autorotation the blade will move well in advance of the radial line until it bottoms out the damper and then the aerodynamic forces back drive the gear box. When you do a flare conservation of angular momentum kicks in and the blades will begin to lead and lag. Push forward cyclic with down collective and the back drive starts and the lead and lag stops.

If lead and lag were to take place during autorotation the driving forces would be like a jackhammer and would make things very unpleasant on the way down.

To: 212man

Periodic drag forces do not cause lead and lag. Please read above. That theory went out the window in the early fifties..

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The Cat

[This message has been edited by Lu Zuckerman (edited 15 January 2001).]

rotorque

15th Jan 2001, 12:56

Lu,

I may have mentioned this one before. Your right in saying that they don't address pitch coupling when they teach basic aerodynamicsm mainly because they use the Bell systems (or similar) to describe dissymetry of lift. On the bells, as you have described many times in the R22 thread, the pitch horns and pitch change rods are placed literally 90 degrees from the blade cord. As the blades flap we do not get a coupling as the position of the horn bearings allows for swivelling in this exact spot.
The reason that the blades do not continue to rise and fall with extremes is due to 'Flapping to equality' (note the term flapping). When the blade flaps up it experiences increased induced flow, which reduces the blades angle of attack and limits the upward motion of the blade, and vica verca on the retreating side.
If you mention a delta 3 hinge or something like it then pitch coupling as I understand it will be the governing factor in limiting flapping motions. But you refer to pitch coupling as the main force.
Once again I should mention that that is how I understand it, If I don't understand your meaning of the words 'pitch coupling' then you may have to go back to the basics for me so I can start thinking along the same lines.

Sorry Helimutt. I thought it prudent to mention this on your thread as it relates to the end result of leading and lagging and corriolis effect. Actually I am getting as much info out of this as you, the rotor droop comments have been great. I have never even 'seen' a graph on the thing.

Cheers

212man

15th Jan 2001, 15:39

I've seen that film too and agree it's not for the faint hearted!

Lu, obviously the lift stays constant as the blade moves around, but are you sure the drag does too? I'd be suprised if the L/D value stays the same, hence the periodic drag variation.

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Another day in paradise

SPS

15th Jan 2001, 17:09

CD=1/2pV2S.

If V2 varies thru' 360 degrees then drag must also vary.

SPS

15th Jan 2001, 17:23

Small but interesting point...

If torque is the twisting force exerted across the fuselage in reaction to the rotation of the rotor blades (the reason we must employ an anti-torque rotor in a coventional design) then how can torque also turn those blades?

I don't think it can. You cannot have it both ways. Torque is a reaction to rotating the blades, something which is achieved by using engine POWER.

helidrvr

15th Jan 2001, 19:37

SPS,

isn't that why we have a tail rotor?

helimutt

15th Jan 2001, 19:42

To all of the posts here. Many thanks for teaching me so much but how much I'll remember is a different matter altogether. It's surprising how much one little question leads to such debate. I have another if anyone is interested?
Also, I also saw a video once where there was a camera attached to the tip of a blade in flight and the pictures were, once again, very scary. I'm trying to see if I have the video as I'll capture a shot from it and JPEG it to show on here maybe.
bye for now.

212man

15th Jan 2001, 21:41

SPS,
Torque isn't supposed to be a measure of power, as you quite rightly say, it isn't! what it does tell you is the stress being sunjected to the transmission and drive train.

It is simply a measure of how much blade drag (plus tiny bit of mechanical loss) there is. The actual power being applied is:

Power=Tq*Nr*K (where K is a constant for that system)

In answer to your query about V squared, don't forget that the A of A changes thereby changing the Cd and Cl. That's what the flapping to equality is all about otherwise using your logic the lift would assymetric too. however, I don't believe that the Cd varies in unison with Cl, so the drag may well change.

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Another day in paradise

lmlanphere

16th Jan 2001, 04:57

torque is the force produced by the engine which turns the rotor. the reaction of the fuselage opposite the rotor, is just that, a REACTION to torque (not to be confused with torque itself; see Newton's law "for every action, blah, blah...")

Lu Zuckerman

16th Jan 2001, 05:16

To: 212 man

“Lu, obviously the lift stays constant as the blade moves around, but are you sure the drag does too? I'd be surprised if the L/D value stays the same, hence the periodic drag variation”.

This is off the top of my head but I would think that the drag would vary with changes in pitch either as collective or cyclic. I would think that the amount of drag would have to be calculated but I don’t think it could be detected relative to lead and lag. Remember in powered flight the blades are always behind the radial position due to the inertia of the rotor system and the introduction of cyclic in directional flight will cause the blades to lead and lag due to Corriolis forces.

If you visualize the drag and how it varies with blade position relative to the oncoming airstream the advancing blade would have greater drag due to the addition of rotational speed and airspeed and it would cause the blade to lag. However the advancing blade leads. Conversely, the retreating blade with the addition of rotor and air speed would lead but in fact it lags.

I would imagine that if the changes in drag across the disc could be calculated it might be found that these changes in drag have some influence on the leading and lagging but they do not cause lead and lag.

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The Cat

Lu Zuckerman

16th Jan 2001, 05:46

To: Rotorque

When I mentioned that when blade flapping is discussed in teaching helicopter aerodynamics they usually discuss a two-bladed rotor system such as a Bell and they never discuss pitch coupling. This was not meant to apply to a Bell rotorhead. They just never discuss it. In that way, everyone that learns that the advancing blade flaps up and the retreating blade flaps down because of the design or the rotorhead (semi rigid) then they have this picture in their mind. The talk about flapping up and down but aside from not discussing pitch coupling they never tell you how far the blades flap up and down. With the picture of flapping in the students mind they could be lead to believe that the amount of blade movement is quite large and if they took that picture one step further they could visualize the helicopter flying backwards.

To my knowledge they don’t use the term “flapping to equality” in teaching helicopter aerodynamics in the US. If they go deep enough on the subject they will introduce pitch coupling. This occurs when the blades, which are normally flying parallel to the swashplate, rise up and try to fly up and away from the parallel path. When this happens the blade that flaps up will move away from the fixed pitch link and as a result will have pitch extracted. The blade that flaps down will do the opposite in that it will have pitch added thus returning the blade to the commanded tip path. The way you describe flapping to equality is taught as pitch coupling in the US.

If you have been following these threads you will know of the beating I took in discussing gyroscopic precession. It wasn’t until one of the participants on these threads explained to me that the UK method of teaching helicopter aerodynamics recognizes that gyroscopic precession exists but it is an aerodynamic force the causes the disc to tilt. I guess we will have to make the same adjustment in speaking of flapping to equality and pitch coupling.

I will repeat what I had stated in other threads. Pitch coupling takes place when the blade flaps up or down relative to the commanded position but only if the pitch horn-pitch link attach points are above or below the cone hinge on the Robbie or the flapping hinge on any other helicopter. The same is true for the Bell with its’ 90-degree pitch horn if the connect point is above or below the teeter hinge. If the points are coincident with each other there is no pitch coupling

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The Cat

rotorque

16th Jan 2001, 13:51

OK, now I'm with ya.

The term 'flapping to equality' is used here in Australia and it refers to an aerodynamic reaction not a physical change caused by the pitch horm position, although 'now' I have the added advantage of knowing that the reaction is a mix of both.

Thanks.

SPS

17th Jan 2001, 12:51

Correction -

CD1/2pv2S.

Same deal, no mistyping.

JoePilot

17th Jan 2001, 21:29

BTW - I've resisted pointing out on other threads that its a pretty strange engineer who doesn't understand Newton's 'Three Laws',
and does't understand the terms: Centripetal/fugal force, mass, weight, force, vectors, fatigue, flapping... etc

I think characters like 212, grey, r.toque, offshore (even young(now modest) RW1) are impressively restrained when dealing with this armchair bluffing amateur.

(I'm also fairly immpressed by HIS learning rate ... if you read HIS history of responses he manages to learn just fast enough not to get humiliated by not understanding the subject, pretty close sometimes though...)

Lu Zuckerman

17th Jan 2001, 22:09

To: Joe Pilot

First of all I would like to congratulate you for your marvelous restraint since coming back on the forum. However I wish you would use that restraint in describing my shortcomings relative to the following;

“BTW - I've resisted pointing out on other threads that its a pretty strange engineer who doesn't understand Newton's 'Three Laws',
and doesn’t understand the terms: Centripetal/fugal force, mass, weight, force, vectors, fatigue, flapping... etc”.

Please explain point by point where I said or implied any thing that was incorrect. When I use the word incorrect I mean by technical standards and not by what you feel is right.

------------------
The Cat

SPS

18th Jan 2001, 14:56

Keeping it brief but putting in the necessary for those that read only.

The way ustd it is that if V2 increases (advancing blade) then the drag associated with the resulting increase in lift(I prefer thrust)production (relying on CL1/2pV2S) must also increase as V2 is an essential element of that lift production. Drag increases exponentially with increase in lift (linked to increasing V2) until stall.

(I have left out increasing collective pitch with speed and many other factors to limit the post and I am ready for bed!)

The reverse must be true of the retreating blade, drag is decreasing exponentially but at a differnt rate because of the (now) different direction of incoming airflow in relation to it (ie reversed).

Drag on the advancing blade increases as drag decreases for the retreating blade as speed increases all the way up to VNE
but this does NOT happen at a proportional rate (advancing against retreating) due to the incoming airflow being in one direction only. By the time VNE is reached AB drag has increased by a far greater value than RB drag has decreased. This is a function of the exponential drag curve being applied to the blade from differing 'start'points on that curve as incoming airflow changes direction. So drag is constantly varying across the disc as speed is changed, whether speed is increased or decreased.

I beleive (although I cannot prove it yet)that this effect is at the root of rotor vibration, in its simplest form portrayed as
two beats per rev. on a semi rigid head such as the (dare say it) Robinson, 206, etc. etc.

Drag is also constantly varying across the disc in still air as invariably some cyclic input is used (if only to neutralise TR drift/roll)and therefore variation in pitch and angle of attack will result in variation in drag values but it ALSO changes
by a vast amount in different areas (or for different blade postions if you like) as speed is increased OR decreased as is shown above.

No, I don't beleive that drag is constant across the disc !

As far as torque goes, that is often misquoted. We generally understand any force to act in only one direction, no matter what that force may be.

"Torque reaction" has been corrupted for a long time. It should read
"Torque - Reaction" or more clearly "Torque,
A reaction"

The blades are moved thru' the air by power (in powered flight) The reaction (Newtons 3rd) is torque. You cannot have torque turning the blades in one direction and turning the fuselage in the opposite.
Power is the action and torque is the equal and opposite reaction,its as plain as that.

And yes, that is why we need a TR or anti - torque rotor but I thought I'd covered that earlier and I doubt if you'd read me fully on it Helidrvr and that's just fine, I often do that too!

212, smack on the button to my thinking..
This is another cause of the misunderstanding of 'torque'. Because we have a 'torquemeter' provided in many helis. it is often thought its reading indicates power in use. NO! it does as you say, shows the amount of torque exerted on the transmission/mast/fuselage etc. right into the air itself and it therefore reads the OPPOSITE of power - Torque. Anyway enough torque, its becoming a drag.... G'nite chaps!

SPS, outside the square on a shrinking oblate spheriod in an expanding universe.

offshoreigor

18th Jan 2001, 15:42

JoePilot:

You hit the nail on the head. If someone wants to dig a hole, why not let them. I agree there have been some great responses here and the amateurs know who they are.

Cheers, OffshoreIgor http://www.pprune.org/ubb/NonCGI/eek.gif

PS, I'm still trying to figure out why anyone would confuse Lead/Lag (fully articulated) with simple flapping. I'll wait for the experts next boola boola burst.

[email protected]

19th Jan 2001, 00:35

Lu, reference your last post -

"This occurs when the blades, which are normally flying parallel to the swashplate, rise up and try to fly up and away from the parallel path. When this happens the blade that flaps up will move away from the fixed pitch link and as a result will have pitch extracted. The blade that flaps down will do the opposite in that it will have pitch added thus returning the blade to the commanded tip path."

The blades cannot flap away from the swashplate as they are connected to it by the pitch operating arms(POA) which are destined to always follow the tilt of said swashplate. For the blades to do anything other than mimic the swashplate tilt, the POAs would have to lengthen which is not a good thing to happen in the air.

The swashplate tilts - the pitch operating arms follow the swashplate and change the pitch cyclicly as they complete their journey around the rotor mast. The change in pitch changes the lift on the blade as CL is dependent on AoA (mostly) and makes the blades flap up or down accordingly. As a blade flaps up the C of G gets closer to the centre of rotation (rotor hub) and the blade tries to speed up due to conservation of angular momentum. Countering this is an increase in drag caused by an increase of pitch and therefore AoA (Cd is also affected by AoA).

As lift changes so must drag as they are dependent on the same formula.

On the subject of static droop - I am led to believe it was used to aid power sharing on twin-engined helicopters with hydro-mechanical governors before FADEC came along. As previously mentioned the static droop slope of Nr against power can be adjusted. If at a specific power setting (or throttle angle) the governor tries to set a specific Nr, then a twin engine installation where the governors are not interconnected, can be encouraged to share the driving load equally.
If you tried this with governors set to give constant Nr despite the power setting, at a given Nr one engine could be working it's t*ts off while the other sat at flight idle as both governors were achieving their desired Nr but with different engine outputs.

Grey Area

19th Jan 2001, 01:22

On the subject of Torque.

Often torque and power are mistakenly imagined as the same thing.

Think of a spanner, you apply a force to one end which is tranlated into a TORQUE, which is simply the magnitude of the force you apply multiplied by the distance of the applied force from the fulcrum (or the centre of rotation although it does not HAVE to rotate). In the olden days it was foot pounds, now newton metres but could be hundredweight cubits, it doesn't matter! It could even be two forces acting as a couple about a fulcrum.

The first thing to bear in mind is that you could measure the force applied at the tail, the rotor blade tips, the gearbox coupling; all would allow you to measure a torque which, ignoring losses, should be the same. You can therefore say that the engines exert a torque at the rotor tips AND the tail boom. Without an anti-torque system the torque couple would cause the fuselage and rotors to turn in opposition.

Second point of academic interest only. As 212man stated, one way power can be described in a rotating system is as a function of torque and Nr. As you increase rotational speed at a fixed power setting so the torque drops, which is why drive shafts always spin so fast, the faster the spin the lighter the shaft (up to a limit). But if you apply full power at low NR you can seriously over torque as we all know (and a collegue of mine can tell you all about that!)

Lu Zuckerman

19th Jan 2001, 02:16

To: Crab

What you seem to forget is that when the blade flaps up or down it is attached to the pitchlink and the pitchlink is fixed by its’ attachment to the swashplate. The blade however is free to feather in relation to the pitchlink. When the blade flaps up, the fixed pitchlink will cause the blade to feather and decrease pitch. The opposite is true for the downward flapping blade. It moves down on the fixed pitch link and pitch is added. This is pitch coupling. Here is how it is explained in the FAA Rotorcraft Flying handbook when they were addressing dissymmetry of lift. “If this condition were allowed to exist, a helicopter with a counterclockwise main rotor blade rotation would roll to the left because of the difference in lift. In reality, the main rotor blades flap and feather automatically to equalize the lift across the rotor disc”. It works exactly the same on a tail rotor. The mechanism that allows this to happen on the tail rotor is a delta hinge. The delta effect also works on the main rotor but only when the pitchlink / pitch horn connect point is above or below the flapping hinge or the cone hinges and teeter hinges on the Robinson or, the teeter hinge on a Bell. If the points are coincident with each other when the blade flaps the blade will pivot about the flapping hinge and the pitchlink/pitch horn attach point and there will be no pitch coupling. This would be extremely rare for these points on all of the blades to be aligned in flight.

I don’t dispute what you said about the drag induced by pitch change and, that the drag is different on the advancing and retreating side. What I said previously is that the drag forces have minimal if no effect on the Coriolis forces that cause lead and lag.

The advancing blade has a higher drag than the retreating side but the advancing blade leads and the retreating blade lags which is in the opposite direction of the respective drag forces.

------------------
The Cat

[This message has been edited by Lu Zuckerman (edited 18 January 2001).]

[email protected]

19th Jan 2001, 12:50

Lu,

The pitch-flap couple you describe is the Delta 3 hinge effect discussed on previous threads which, as you correctly stated, can only occur if the pitch change horn is not in line with the flapping hinge. Most examples of this effect are found on tail rotors where the flapping limits must be constrained to keep the blades clear of the tail structure and droop stops or flapping restrainers are undesireable alternatives because they add weight.
Back to the main rotor; the reason the blades rate of flapping reduces is due to a reduction in cyclic pitch as the POAs follow the control orbit and aerodynamic damping - as a blade flaps up, it's relative airflow changes direction and gradually returns the angle of attack towards it's starting value.
Tilt the swashplate forwards (towards the nose) the cyclic pitch measured at the 3 and 9 0'clock positions is the same - the minimun pitch is 12 0'clock and the maximum is 6 o'clock.
As a blade moves from 3 o'clock towards the nose it begin to see a reduction in pitch (POAs follow swashplate) and therefore AoA and starts to move (flap) downwards. The rate of pitch change is not uniform and the maximum rate (vertical movement of the POA for a given angular displacement) of pitch change is at the 12 o'clock position.
At this point the rate of flapping down is at it's maximum as well and although the cyclic pitch from 12 to 9 0'clock is increasing again, it is still less than the 3 and 9 0'clock settings and the blade continues to flap down until it reaches it's low point at 9 0'clock. The blade is slowed in it's rate of flapping in the manner described above but as the blade is flapping down the relative airflow is coming from a greater angle so the AoA is increasing.
The 90 degree disparity between swashplate tilt and blade low point is Phase Lag - the way the blades AoA is reduced or increased by flapping is Flapping to Equality and exactly the same phenomenon is the reason for Flapback (equalising the changes in lift caused by speed differences between advancing and retreating blades in forward flight) and Inflow Roll (equalising changes in lift between the front and rear of the disc caused by disc tilt and the resulting differences in relative airflow).

That hurt my brain so I'm off to work for a rest!

Lu Zuckerman

19th Jan 2001, 16:58

To: Crab

First of all we have to get on the same piece of paper. In this discussion I will not address the Lynx or the Robinson because of their respective offsets (Lynx 15-degrees and Robinson 18-degrees). I will not address swashplate position; instead I will address cyclic stick displacement, as this is uniform on all helicopters where swashplate displacement is not.

First lets address the Bell. If the blades are displaced in line with the lateral axis and the cyclic stick is moved forward the blade at the 3:00 position will be at the minimum pitch angle. The blade at the 9:00 position will be at the highest pitch angle. Gyroscopic precession will tilt the disc down over the nose and up over the tail. (If you don’t agree with gyroscopic precession you can use your own terminology). If the blades were disposed over the longitudinal axis and the cyclic stick was moved forward or backward the blades would not change pitch. If the blades were disposed over the longitudinal axis and the stick was moved to the right the blade at 12:00 would be at its’ highest pitch and the blade at 6:00 would be at its’ lowest pitch. The opposite would be true if the stick were moved left.

Now for clarity’s sake I will use a four-blade rotor system turning counterclockwise in this illustration. With the blades disposed over the lateral and longitudinal axes of the helicopter and the cyclic was moved forward the blade at 3:00 would be at its lowest pitch and the blade at 9:00 will be at its’ highest pitch. The blades at 6:00 and 12:00 will reflect the collective pitch setting. If the cyclic were moved to the right, the blade at 12:00 would be at its’ highest pitch and the blade at 6:00 would be at its’ lowest pitch. The blades at 3:00 and 9:00 would reflect the collective pitch setting.

All of this is allowed by the lead angle of the pitch horn and the tilting of the swashplate in respect to the direction in which the cyclic stick is displaced.

With all of this in mind, you are free to go back to the “drawing board” and rethink your post.

------------------
The Cat

JoePilot

20th Jan 2001, 00:13

Lu writes (to Crab):
"With all of this in mind, you are free to go back to the “drawing board” and rethink your post."
Very condescending to distinguished (and modest IMO) Crab ... it's fairly clear that he just accidentally shifted his points of reference accidentally (twice)… but still has a better understanding than U Lu..

..AND excuse me for suggesting but I think Lu YOU might benefit from a bit of a rethink.

Since you do not seem to understand the nature of helicopter rotors. Hiding behind the term Gyroscopic Precession (GP) is becoming boring...
Not understanding Flapping to Equality is pretty odd (not that it actually occurs much in practice)...

Try this for an unusual (but simpler) approach:
(For S&L cruise, teetering head)
Terms;
Disc Attitude (angle of Tip Path Plane Axis* from vertical) I'll call X

X=invTan(DRAG/WEIGHT)

(….because DRAG=Tsin(X) and WEIGHT=Tcos(X), where T=TotalRotorThrust)

So the Plane of Rotation for a given cruise is predetermined!...*

Now all you have to do is arrange for that to continue being the case...

For this to continue the 'Lift' generated by a blade through the Cycle must remain the same through one half rotation as the other half (any of the infinite possible halves will do).
IF it does not THEN the consequent response of the blades to Flap to Equality* (Lift actually still remaining the same* all the way around the disc) will RESULT in a change of Attitude* ... This change will be opposed by the PILOT, who varies the asymmetry in Pitch (Cyclic Pitch) such that the attitude does not need to change*.

So really it IS actually the pilot, using the feedback of Attitude, who eliminates dissymmetry of lift*, by running his Swash Plate at whatever attitude is required.

It would make NO material difference if Delta3 Pitch Coupling existed or not, the pitch the blades ran would still have to be the same* with it or without it it's just that the pilot would have to hold his stick in a different place to achieve this… in the same way he would if the passengers walked around (assuming no drag change from fuselage attitude change) …

Pilots generally do not mind moving the stick. They often actually enjoy it … which I suppose is why 'Auto' Pilots (airline pilots) love to learn to fly helicopters … if they didn't already…

Good morning.

* = Things Lu won't understand….

Lu: "Here is how it is explained in the FAA Rotorcraft Flying handbook when they were addressing dissymmetry of lift. “If this condition were allowed to exist, a helicopter with a counterclockwise main rotor blade rotation would roll to the left because of the difference in lift. In reality, the main rotor blades flap and feather automatically to equalize the lift across the rotor disc”. " - trying to explain how Pitch coupling is responsible for eliminating Diss-o-Lift - VERY ODD ... now Stability ...that's another story…

JoePilot

20th Jan 2001, 00:38

Actually to be fair I think this forum is really helping Lu learn about helicopters - and he's starting to make some sense, it's just amazing that he keeps his crediblity (retrospectively) throughout his learning process ... hats off to all

How?

[email protected]

20th Jan 2001, 00:59

Lu,

I agree with all you say regarding the blade pitch response to cyclic position - I should have stated that I was using as simple an example as possible where the pitch change arms would be positioned 90 degrees ahead of the blade pitching axis on a Bell type 2 bladed rotor. In my example, to get the blade low at 12 'o'clock when you move the cyclic forward, the swash plate is tilted to the 3 o'clock as the blade will reach it's low position 90 degrees after the minimum pitch position (I say due to flapping - you say due to precession). We are both on the same piece of paper and if you go back and re-read my post with that in mind you will see we agree.

I was trying to clarify Flapping to Equality as used by Brit Mil and many others as it is not the same as the pitch-flap coupling you described which we regard as the Delta 3 hinge effect. Your previous post stated that the blade was free to feather in relation to the pitch-link which is not the case - it would be impossible to control the aircraft if the blades could increase or decrease pitch except as controlled by the pitch operating arms which are fixed to the swashplate.

You might not like the idea of aerodynamic forces and prefer precession but if it's good enough for the Empire Test Pilot's School (ETPS) who's notes I have read, then it's good enough for me!

Lu Zuckerman

20th Jan 2001, 20:03

To: Crab

Your illustration of the Bell swashplate Vs blade tilt is wrong. When you push the cyclic forward on most Bell helicopters the swashplate will tilt down at the 12:00 position with the blades disposed over the 3:00 and 9:00 positions. With the blade in this position the pitch horns are at the 12:00 and 6:00 positions. That means that the blades have the maximum pitch change relative to their collective pitch angles. The advancing blade will have less pitch and the retreating blade has more pitch and some “unknown force” will cause the blade disc to tilt down over the nose.

In stating that the blade is free to feather a better choice of words could have been is that the blades can be made to feather. This feathering can be accomplished by moving the pitchlink in relation to the blade or, by moving the blade in relation to the pitch link. This is what happens when the blade flaps up or down in relation to the parallel paths of the disc and the swashplate resulting in pitch coupling.

------------------
The Cat

Grey Area

20th Jan 2001, 23:40

:rolleyes: Is there a parrot around here, or is it an echo from before Christmas? :rolleyes:

lmlanphere

21st Jan 2001, 02:19

sps, the lose definition of torque is: a force which tends to produce rotation or torsion. That definition doesn't include anything about a "reaction". if you wrap your hand aroud a shaft and rotate it (don't try this at home) you are exerting the force (torque) directly to the shaft. maybe a more accurate name for the tail rotor would be anti-torque reaction system. maybe not.

Lu Zuckerman

21st Jan 2001, 04:50

To: Joe Pilot et al.

The reason I am starting to make sense is summed up in the story of the young man that thought his father was so stupid, and when the boy reached 21 years of age, he was amazed at how smart his father got in six years.

It must be understood that the major reason for so much antagonism from UK and OZ pilots and their not agreeing with my input on these threads is due to the fact that your training regarding flight theory is totally different from how it is taught in the USA. With that said, I find it difficult to understand why so many UK and OZ pilots quote books and theories postulated by people like Ray Prouty and other American text book writers. How do you reconcile the differences in aerodynamic and physical laws as taught in the US and the UK?

Regarding my quoting the FAA Handbook, the quoted statement should have been preceded with; …The relative wind encountered by the advancing blade is increased by the forward speed of the helicopter while the relative wind speed acting on the retreating blade is reduced by the helicopters forward speed. Therefore, as a result of the relative wind speed, the advancing blade side of the rotor produces more lift than the retreating blade side. This situation is described as disymmetry of lift. Then comes the paragraph I quoted above. The automatic feathering they alluded to is pitch coupling.

If you have an argument, take it up with the FAA because I don’t agree with their explanation. I believe disymmetry of lift, if it exists, results in blow back. If you look at the rotor system that has disymmetry of lift across the rotor disc it appears that it is basically the same as that of retreating blade stall but a much milder version. That’s what I think. The FAA states that the helicopter would roll left under the stated conditions. That is the first stage of retreating blade stall. If the left roll were not countered the disc would blowback, and you might ask, what would cause the disc to blow back. It will blow back for the same reason the disc blows back with retreating blade stall. The culprit (are you ready for this?) is gyroscopic precession.

One of the reasons I disagree with the FAA is that they address individual subjects in an isolated manner much like the way helicopter aerodynamics is taught. They allude to the fact that the blades are automatically feathered (pitch coupling) but they neglect to discuss one very important point. That point is that in order to get the helicopter to fly through the relative wind that will create disymmetry of lift you have to move the cyclic forward which will decrease the pitch on the advancing blade and increase the pitch on the retreating blade thus tilting the disc and at the same time result in eliminating disymmetry of lift. There are other perturbing forces out there that will cause a helicopter to roll right or left one of, which is transverse, flow effect.

Let the race begin.

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The Cat

[This message has been edited by Lu Zuckerman (edited 21 January 2001).]

[This message has been edited by Lu Zuckerman (edited 21 January 2001).]

SPS

21st Jan 2001, 14:36

I have read it. What of it? It could be saying many things that might have some bearing on the thread on flapback/forward, and then again it could have no relevance at all with regard to my actual question.

I know why the disc does not flap forward as much as it flaps back below 80% of BRC speed.
(Well, I think I do!) Part of the reason
that the flapping thread was started was to see what others have to say on the matter, and it was done with an open mind.

Gyroscopic precession has a part to play in the process but I do not beleive it to be a panacea for all P of F anomilies that initially seem to be unanswerable without reaching for it.

As I said to you before, US training may well
rely far too heavily on Gyroscopic Precession whilst UK training may not use it enough. I have a feeling that the truth lies somewhere in the middle. You didn't choose to reply to me on that point, your choice of course.

In any event, if you have something to add to what has been said on the thread then I'd be really pleased to hear it and have the opportunity to discuss it with you, OK?

SPS, watching the water go down the plughole the opposite way whilst my head still goes the same way as it did in the N. hemi after a few beers...?

[email protected]

21st Jan 2001, 17:39

Lu,

In an attempt at clarification I muddied the waters - my simple 2 bladed model had double correction for phase lag, both swash plate tilt to the right to produce blade low at the front and a 90 degree advance angle on the pitch operating arms. One or other of these corrections would have been alright and the aim was to agree that the blade reaches it's high or low position 90 degrees after the maximum or minimum cyclic pitch position.
Most real helicopters use a combination of jack positioning and advance angle of the POAs to achieve the obviously desireable state where forward cyclic makes the blades flap down at the front and up at the back, thereby angling the Total Reaction forwards and accelerating the helicopter.
If the FAA thinks that differential airspeeds on the advancing and retreating blades makes the aircraft roll left then good luck to anyone flying a helicopter certified by them. The true result as we all know is flapback (blowback).
The roll towards the advancing blade (to the right on a counter-clockwise rotor)is a result of disc tilt causing a reduction in the inflow angle at the front of the disc. The increase in AoA gives more lift and as the effect starts at the 3 o'clock position, has it's maximum at 12 o'clock and reduces by the 9 o'clock the disc flaps to the right (90 degree phase lag again). We call it Inflow Roll and you call it Transverse Flow.

JoePilot

21st Jan 2001, 19:56

Lu WROTE: "The reason I am starting to make sense is summed up in the story of the young man that thought his father was so stupid, and when the boy reached 21 years of age, he was amazed at how smart his father got in six years."
- NONSENSE you are just learning, without anyone realising how little you knew...

- The Pilots job in setting a Cyclic dissymetry to counter aerodynamic dissymetries to ensure equal lift around the disc ... you have clearly just understood!

Shame you are taking so long to understand what you mean by Gyroscopic Precession:
GP: More pitch on ONE side will cause the blades of that side to flap (a process) up, the process of flapping up causes the Angle of Attack to be reduced. Result: the 'lift' stays the same but only because the blade is now RISING - this RISING (or Flapping Up) will finnish at the END of the half where there was extra pitch, ie. APPROXIMATELY 90degrees after point of MAX Pitch.
It is very convienient not to have to explain this, so the term 'Gyroscopic Precession' is used since it BLINDLY predicts approximately the same behaviour.
You must try to understand this - otherwise you miss the true 'elegance' of the helicopter.

Lu:"How do you reconcile the differences in aerodynamic and physical laws as taught in the US and the UK?" - Simplification for Americans... of course!

On reading Lu more carefully it is apparent that his use of English is often what causes the misunderstandings.
Allow me to illustrate Lu's ENGLISH being 90Deg 'out of phase':
Lu: "When you push the cyclic forward on most Bell helicopters the swashplate will tilt down at the 12:00 position with the blades disposed over the 3:00 and 9:00 positions. With the blade in this position the pitch horns are at the 12:00 and 6:00 positions. That means that the blades have the maximum pitch change relative to their collective pitch angles. The advancing blade will have less pitch and the retreating blade has more pitch and some "unknown force" will cause the blade disc"

In this paragraph the phrase: "blades have the maximum pitch change" implies that the rate of change of pitch is a max - whereas I'm sure he means the pitch (ammount, quantity or value) is a Maximum (or minimum). Pitch Change implies a rate, the max rate of which obviously occurs at 12&6 o'clock.

I'm sure CRAB is just getting HIS TERMINOLOGY 90deg out from our understanding in the same way, since he does at least appear to UNDERSTAND the process:
A blade at 3 has it's Pitch Link at 12 (approx), therefore to run min blade pitch at 3 the swash plate must be 'tilted down' at 12. The Max Rate of Pitch Change occurrs at 12(blade at 12, pitch link at 9) requiring upward Force on swash plate at 9 hence righthand stick force(not movement, just force) requirement. Right CRAB?

Lu: I guess the simplest way to get accross the concept you call GP is to think of the pitch variations flying the blades to a new plane of rotation - does that help?

(and/or likewise ...preventing changes in plane of rotation)

Lu you are clearly still 'screwed up' over things like 'Blow Back' - to us this is more violent version of 'Flap Back' - do you use that term? Retreating blade stall and Diss of Lift - similar ? - well yes ... but only in-as-much as that IS how a rotor system works!

Lu:"The culprit (are you ready for this?) is gyroscopic precession." - no! Again: A Falling blade (flappING down) will be low at the END of the half in which it is FALLING. .. getting it yet?

CRAB (and other mil-bible die hards)... just a side question : Do the military still teach (wrongly) that recirculation at cliffs results in attitude change TOWARDS the cliff? (worth a (see)new topic(s)... I think)

[This message has been edited by JoePilot (edited 21 January 2001).]

[email protected]

21st Jan 2001, 23:26

Joe,

I was trying to show Lu that although we disagree on the Precession issue, he was arguing correctly about blade low/high being 90 degrees after the min/max pitch position.
I wanted to highlight that you either have to tilt the swashplate 90 degrees early or mount the pitch operating arms 90 degrees ahead of the blade axis. Unfortunately I applied a both to my second post because Lu wanted to refer to cyclic displacement rather than swashplate tilt and rather confused the very thing I was trying to clarify.
I can see you have been a student of the "In your face" school of communication - did you study with RW-1?

Lu Zuckerman

22nd Jan 2001, 05:21

To: Joe Pilot et Al.

First of all when I wrote that bit about the father and son I was visualizing you as the son. In other words sooner or later you would come to realize that the father was correct. As long as you stay in the UK or OZ you will always be right and I will be perceived as an idiot. However if you ever have to come the USA for training say at Sikorsky and they start talking about theory of flight they will have to put you in restraints because based on how you respond to me on this forum you will most likely become violent.

Now I’m going to tell you a story about the life cycle of a helicopter blade or should I better use the term rotational cycle as that would be proper English. What I am about to say is applicable to every counterclockwise rotor system with the exception of a Robinson and a Westland Lynx. It applies also to clockwise rotation blade systems but the clock positions of 9:00 and 3:00 are switched, otherwise everything is the same. All numbers denoting pitch settings are applicable to this description and do not apply to a specific helicopter. Another point, when I use the term maximum pitch change relative to collective I am not addressing rate of change as that is constant as it reflects the angular deflection of the swashplate. There are several other disclaimers one of, which is cyclic pitch change to counter tail rotor propeller effect, is not considered. Another is that the collective pitch once set will not change to compensate for translational lift. We will use a four-blade rotor for simplification.

1) The rotor system is at rest. We measure the basic collective pitch setting at the root of the blade and it is 6-degrees with the cyclic in neutral.

2) The collective range is set at 18-degrees.

3) The rotor is started and the helicopter is brought to a hover. If we could measure the collective pitch setting in a hover it is 18-degrees on all blades as they rotate.

4) The pilot pushes forward cyclic. If we could stop the rotor with the blades disposed over the longitudinal and lateral axes the pitch readings would be 3:00= 12-degrees, 12:00= 18-degrees, 9:00= 24-degrees and 6:00= 18-degrees.

5) At these respective positions 3:00 and 9:00 the blades have the maximum variation relative to the basic collective setting of 18-degrees and rate of change has nothing to do with it.

6) When the blades move from the respective positions of 3:00 and 9:00 the pitch will start to change until the blades are disposed over the longitudinal center line of the helicopter and at that time they will be at the full collective position of 18-degrees. They have increased and decreased pitch respectively.

Now try to visualize this. If what you say is true about the advancing blade flapping up and the retreating blade flapping down due to changing aerodynamic loads then for gods sake please explain how or what forces are involved to tip the disc down over the nose. You say the advancing blade flaps up and the retreating blade flaps down. How much is the flap up and flap down and why won’t the delta three effect cancel this movement. Nobody ever tells you how much, he or she just says it happens and the rest is left up to your imagination.

Another point to ponder, if the advancing blade is increasing pitch between 3:00 and 12:00 and the retreating blade is decreasing pitch between 9:00 and 6:00 how can the aerodynamic forces cause the disc to tilt down over the nose and up over the tail? The way you explain it the disc would rise over the nose and down over the tail if it were pure aerodynamics. Yet, when the pilot pushes forward cyclic the disc tips down over the nose. The next time you fly your helicopter, watch the tip path when you move the cyclic in any direction. If when you address flapping up and down in forward flight and the rotor disc tips back this phenomenon also results from an increased lift on one side of the disc causing the disc to blow back(FAA definition) the blowing back is caused by the gyroscopic precession effect.

In the United States where they respect the laws of physics and the laws as they apply to gyroscopic phenomenon they would look at it in this way. When the pilot first pushed forward cyclic he changed the pitch relationship between the advancing side and the retreating side. This creates a differential of lift due to the pitch differences and this equates to a force that is applied to a spinning rotor. The lift differential is just like the perturbing force on a gyroscope and the application of the greater force is on the left side of the disc. And like all good gyros the response will be 90-degrees later in the direction of rotation. Once the forces are stabilized (at the time the cyclic stopped moving) the perturbing force is removed and since the rotor exhibits the gyroscopic characteristic it will remain rigid in that position unless the pilot moves his cyclic or some aerodynamic generated force perturbs the rotor. In this case read Transverse flow effect or a lift imbalance cause during retreating blade stall or the lift imbalance resulting in blowback as described above.

I have said this many times and it still goes un- noticed. Do not look at the spinning rotor as individual blades. The disc moves as a unit and is a composite of the blades made up of that disc. The only time blades act individually and fly out of track the delta 3 pitch coupling will return the wayward blade.

One last thing, when you all went through pilot training 101 you must have been taught how the instruments on your panel work. How did they address gyroscopic precession as it applied to the specific gyro operated instruments? What you rely on for your compasses, your attitude indicator and your turn indicator and some not mentioned all work using gyros and they all work under exactly the same rules that govern your rotorheads.

I think if you cut back on the trifle and the spotted dick your system might clear up so that you can accept an idea that is foreign to you.

Here is another thought and that is the AH 56 Cheyenne was not controlled directly by the pilot when he moved his cyclic. When he did move the cyclic it actuated the servo and the output of the servo was coupled to a spring. The spring was in turn linked to the swashplate which was in turn linked to a very powerful gyroscope mounted on top of the rotor head. When the pilot made the perturbing force via his servo the gyro responded to its’ maximum deflection 90-degrees later and it was linked to the pitch horns on the blade and they responded 90-degrees later at least they did most of the time. Sometimes they would respond early or late and on two occasions they responded with such fury that the blades impacted the fuselage. On one ship it killed the pilot and on the other it destroyed a wind tunnel at Ames Labs in San Francisco, California.

Guess who designed the blades? Ray Prouty.

------------------
The Cat

[This message has been edited by Lu Zuckerman (edited 22 January 2001).]

[This message has been edited by Lu Zuckerman (edited 22 January 2001).]

Lu Zuckerman

22nd Jan 2001, 21:18

To: Joe Pilot et Al

I know this is a bit self serving but I wanted this to go to the top of the line in order to attract your attention and get a response to the above post/

------------------
The Cat

[email protected]

23rd Jan 2001, 00:30

OK Lu let's start again.

1. Assume that the pitch operating arms are mounted to change the pitch of the blade exactly matching the tilt of the swashplate (no advance angle).
2. Assume that the swashplate tilts in the same direction as the cyclic is moved.
3. Use the same 18 degree collective pitch setting as you did.
4. Push the cyclic forward to set the same pitch settings as you had = 18 degrees at 3:00 and 9:00, 24 degrees at 6:00 and 18 degrees at 12:00.

Happy so far?

5. Start with a blade at 3:00 - as it moves towards 12:00 it sees a change in cyclic pitch that reduces tha AoA so the blade loses lift and begins to flap down.
6. At 12:00 it sees the minimum cyclic pitch setting and so has the least lift - it cannot start to flap up as it has 6 degrees less pitch than when it started to flap down. Therefore it continues to flap down past the 12:00 BUT at a reducing rate because the cyclic pitch is increasing back towards the neutral 18 degree setting at 9:00. THE BLADE IS AT IT'S LOWEST POINT AT 9:00 DUE TO PHASE LAG.
7. For the remaining 180 degrees the opposite happens and this is why maximum/minimum cyclic pitch must be applied 90 degrees before the desired blade high/low point. This is achieved by tilting the swashplate and/or mounting the pitch operating arms ahead of the blade feathering axis. THIS IS NOT A MYTH AND ALL HELICOPTERS (except oddballs like the one you described) USE VARIATIONS OF THIS IDEA.

8. The phase lag is only truly 90 degrees on a teetering head rotor system and the figure reduces as the flapping hinge is moved outboard (hinge offset).

The rate of cyclic pitch change is not constant although the rate of angular displacement is :

Draw a clock face with a horizontal line underneath it so 6:00 is touching the line. Draw vertical lines down from 1:00, 2:00, 3:00 and 4:00 to the horizontal line. From 12:00 to 1:00 is 30 degrees and so is 1:00 to 2:00 and 2:00 to 3:00 and 3:00 to 4:00. This is constant angular displacement.

Now look at the distances between each of the vertical lines where they cross the horizontal line - they are not equal. The distance between 12:00 and 1:00 is greater than 1:00 to 2:00 which is greater than 2:00 to 3:00. The pitch change arm moves up and down at different amounts depending on which 30 degree segment of the circle it is in and if plotted would produce a perfect sine wave.
On a swashplate system (known as a control orbit) the maximum rate of pitch change occurs during the 30 degrees arc either side of the maximum/minimum pitch setting which is surely proof enough that the blade cannot start to flap up again at 12:00 in the above example.

The only time the precession causes the 90 degree phase change is in a theoretical model of a rotor system in a vacuum where aerodynamic forces are completely discounted. This model is used to illustrate what subsequently happens when aerodynamic forces are introduced - ie the real world and not theoretical physics. The 2 effects have the same end state but only the aerodynamic effects are generally considered since we don't fly helicopters in space!

Please read this carefully Lu and then go and read Prouty/Padfield/Gessow and Myers. Then come back and tell me if anything they say is different to the above (it's not).

Or.. look at a helicopter rotor head and examine carefully where the swash plate tilts when you move the cyclic and also how far angularly displaced ahead of the blade are the pitch operating arms.

Sorry to go on but this does need sorting.

If you still disagree then I will e-mail you extracts from appropriate publications to show I am not making this up or blindly quoting Brit-Mil dogma.

helisphere

23rd Jan 2001, 00:54

How is disymmetrty of lift compensated for?

I know most text books I've read describe the advancing blade flapping up due to the increase in airspeed. The flapping up then reduces angle of attack because the now upward direction of the blade changes the relative wind. And the retreating blade basically does the opposite. While I argee that by itself this theory of flapping to equality is valid I don't see it being valid in the question of dysimmetry of lift with a helicopter in forward flight.

First of all the book says that the advancing blade flaps up and the retreating blade flaps down. If we simply look at a helicopter in forward flight we instantly see that this cannot be true. If the disc is tilted forward, and, as someone else stated earlier, the tip path plane remains at a constant angle to maintain thrust and lift the same, then how can the advancing blade be moving up? With the disc tilted forward the advancing blade is moving DOWNWARD and the retreating blade is moving UPWARD. This is the opposite of what the book says. Forward cyclic is what takes angle of attack away from the advancing blade and gives more to the retreating.

If you go back to the days of the gyro planes you will see where the theory of flapping to equality compensating for dysymmetry of lift came from. Juan de la cierva's gyros had no cyclic or any type of attitude control in their rotor systems, they also did not have any flapping hinges. So he came up with the idea of putting a flapping hinge on each blade theorizing that the blades would flap to equality thereby reducing an unwanted rolling moment (disymmetry of lift). It seems to work, because there was no cyclic in these designs. I think he did however design the flapping hinge at an angle to produce delta 3 pitch coupling. Think about the disc orientation on a gyro though. In forward flight it is tilted back and you can see that the advancing blade is flapping up and the retreating blade is descending, just like the book says. Maybe whoever wrote the book just figured this must apply to helicopters too. And everyone else who wrote a book just followed suit.

[This message has been edited by helisphere (edited 22 January 2001).]

RW-1

23rd Jan 2001, 01:13

>>(even young(now modest) RW1) <<

Hey JoePilot,

1 outta 2, not bad! (young? yep, that would be me hehe) :)

(hey, visit my site and leave me an email, I'd like to chat private. Email addy is safe with me.)

Lu Zuckerman

23rd Jan 2001, 02:45

To: Helisphere

You have finally discovered the secret. Helicopter aerodynamics is taught in the following way.

1) They discuss Bernoulli’s theorem and then address how a venturi works.
2) They apply this to the operation of an airfoil of a fixed wing aircraft and explain how it generates lift
3) They in turn apply this theory to a rotating airfoil and they use the autogyro as an example
4) Without shifting gears they start talking about helicopters and apply the logic of the autogyro to the helicopter and that is where flapping to equality comes from.

I brought this subject up in the Robinson certification thread.

Thanks for bringing it up on this thread.

To: Crab

It looks like we are about the rekindle the war of 1812. If I remember my history the United States defeated the British. Just a bit of humor to get us started.

If I read you correctly you are stating that the lead angle of the pitch horn in your illustration is zero. If I do understand this premise it would appear that everything you say after that point is contrived to fit the basic premise. Also your diagram of the circle and the horizontal line with the vertical lines establishing the angular distance traveled (or in other words the amount of pitch change) is also flawed. If everything in the system is rotating about a fixed point then the angular displacements must be referenced to the point of rotation. That way, you can divide the circle into equal parts and the distance between each two points is equal to the distance between any other two points.

I have in front of me the Sikorsky Blue book. This book is issued to anyone that attends a technical school at Sikorsky whether they are a pilot or a mechanic. On page thirty of the blue book there is an illustration of blade pitch change as the blades rotate 360-degrtees. Although the figures are different from those quoted in my post the angular change of the blade is consistent for each degree of travel around the circular path.

I don’t know if you are familiar with a constant pressure variable delivery hydraulic pump but it too has a swashplate. As the internal mechanism rotates the swashplate action causes pistons to move in and out of a cylinder that resembles the cylinder in a Webley (spelling) revolver. The movement is constant and can be diagrammed just like a sine wave. If the pistons moved in the way you diagrammed the movement by drawing vertical lines from a base line the pump would tear itself apart and if it didn’t your hydraulic system would have to incorporate a pulse damper or, an accumulator. The pump described does not require an accumulator because of the smoothness of the output. Using this same analogy on the rotorhead, if the angular pitch change were not consistent with the degrees of rotation then you would have a constant medium to high frequency vertical beat in your rotor system.

Also, in your illustration of the 0-degree pitch horn the blade would have to be directly over the lowest part of the swashplate (down over the nose) in order to get the maximum pitch change which would place the blade at its’ lowest pitch relative to collective. Assuming it is possible to do this if using gyroscopic theory the disc would tilt to the left when you pushed the cyclic forward. Which way would it tilt using your theory?

The rules are the same for a bell with a 90-degree pitch horn and a Sikorsky using a 45-degree pitch horn and a 45-degree offset of the swashplate. In either case the lead relative to the selected direction of flight is 90-degrees. You stated in a past posting that the Lynx was like the Robinson in that in rigging the helicopter the blades were offset by 15-degrees on the Lynx and I stated that the Robinson blades were offset by 18-degrees. The result of this offset is that the blade will have maximum response 90-degrees later in rotation. My theory about the Robinson is yet to be proved but you had indicated in a previous post that the Lynx without the Auto Control System would roll to the left or 90-degrees later in rotation from the offset position.

In every case, the figure 90-degrees keeps popping up. Why does this figure keep coming up when discussing rotary wing flight? The answer is simple if you are on this side of the pond. It pertains to gyroscopic precession.

Regarding phase lag, it is 90-degrees whether you are addressing a Bell or a Sikorsky. As stated above the Bell pitch horn leads the blade by 90-degrees and on the Sikorsky it leads by 45-degrees and the swashplate tips down 45-degrees ahead of the selected direction of flight. Most helicopters excepting Boeing CH-47s and CH-46s follow this method of cyclic control input. But even on the Boeing helicopters the rotor placement is governed by gyroscopic precession.

Regarding your illustration about the only time that precession would play a part in the displacement of the rotorhead would be in a vacuum. I don't think that would work as the only thing you would do is change the angular setting of the blades but since there is no atmosphere you cant generate the lifting forces that would cause the disc to tilt due to precession.

In your last paragraph you asked me to look at a rotorhead and swashplate and see the tilting of the swashplate relative to where the blade pitch horns are. Assuming a Bell rotor and it is disposed across the lateral axis and the cyclic stick were displaced forward on the center of the rigged axis then the swash plate would move down forward and up aft. (Not on all Bell two blade systems as on some the swash plate moves opposite but the pitch horn is on the rear of the blades so the angular displacement related to blade position would be the same. On a Sikorsky as described above when the cyclic is pushed forward the swashplate will tip down 45-degrees ahead of the longitudinal axis and up 45-degrees ahead of the lateral axis. The blades if disposed over the lateral axis will have minimum pitch on 3:00 and maximum pitch at 9:00 relative to the basic collective pitch setting.

Over.

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The Cat

[This message has been edited by Lu Zuckerman (edited 22 January 2001).]

helisphere

23rd Jan 2001, 05:00

Question:

Why does the Lynx have a 75 degree lead on the pitch horn instead of 90? I'm pretty sure I know, but I wonder if there is more to it.

At the risk of taking another name calling from Lu. this is my understanding:

In a two blade teetering rotor, the flapping axis is the same as the rotational axis. Therefore the natural frequency of the blade rotationally is equal to the natural frequency of the blade in flapping. This will put phase lag right at 90 deg. However, in a rotor with offset hinges, or in a rigid rotor or bearingless type, the flapping length of the blade is shorter than the rotational length. This makes the natural frequency in flapping higher than the one for rotation. This means that the blade will flap to its highest or lowest point in less than 90 degrees from the point of highest or lowest blade pitch, in other words the phase lag will be less than 90 deg because the blade will flap faster with a higher natural frequency. Lu mentioned earlier that we need to look at the disk as a whole and not so much at the individual blades. However, when cyclic is input to the rotor, it is the tilting of the rotor thrust vector in relation to the rotor mast that produces a rolling moment. This means that there is a difference in the angle of the tip path plane in relation to the hub and mast IE flapping, and we must consider the higher frequency in flapping when figuring phase lag. Now, this is why I understand that all rotor systems with hinge offset have a phase lag something less than 90 deg although with small hinge offset the amount less than 90 is also small. This does not take into account any affect of delta 3 on phase lag or any other theory for that matter, which is why I post the question. Why does the lynx have 75 deg lead and (according to Lu) all other helicopters have 90 (except Robinson)? I happen to know that the sikorsky S-69 did not have 90 deg phase lag, it was quite small like 60 or 70 some degrees because the blades were so rigid. I know this because I have read the test data. I also know that they adjusted the phase angle at high forward speed but not to alter control reponse but to unload the retreating blades to avoid stall and reduce drag as they were flying at very high advance ratios and most or all of the retreating blade was flying backwards.

[This message has been edited by helisphere (edited 23 January 2001).]

Lu Zuckerman

23rd Jan 2001, 05:58

To: Helisphere

I know absolutely nothing about the S-69 ABS helicopter however I would think that the control system must have been extremely complex to change the pitch of the retreating blade to unload it. But, there are some really tricky mechanical devices in this world.

In my post to Crab I used the Bell and Sikorsky rotor systems as examples, indicating one had a 90-degree lead on the pitch horn, and the other had a 45-degree pitch horn and the swashplate was offset by 45-degrees. This gives a total lead of 90-degrees relative to the direction of cyclic movement. The French helicopters are a bit different but then the French are a bit different.

On the A Star and quite possibly on the other Eurocopters helicopters the pitch horn leads the blade by 60-degrees and the swashplate is offset by 30-degrees which gives a total phase angle of 90-degrees. Whenever you see a helicopter that has pitch horns that do not lead by 90-degrees or, 45-degrees you have to check the layout of the swashplate. I would hazard a guess that the total of the swashplate offset and the pitch horn lead will come out to 90-degrees.

Regarding why the Lynx is the way it is I will yield to Crab.

------------------
The Cat

helisphere

23rd Jan 2001, 08:46

I need to rephrase my question. The Lynx being four bladed probably does not have a PITCH HORN leading 75 degrees or it would be very close to running into the next blade. Either way, what I meant to ask was: Why does the lynx have a PHASE ANGLE lead of 75 degrees as was previously said on the forum?

[This message has been edited by helisphere (edited 23 January 2001).]

[email protected]

23rd Jan 2001, 11:20

Lu,
The only mistake I made in my post was regarding where the maximum rate of pitch change occurs. It is in the 30 degree arc either side of the neutral position - the reason the maximum rate of flapping happens at the position of maximum/minimum pitch is because the vertical movement of the pitch operating arms is at it's least 30 degrees either side of this position. This means that the blade spends 60 degrees of it's rotation at or near maximum pitch and 60 at or near minimum pitch. Angle of attack determines lift and therefore flapping so the max rate of flapping is still at max and min pitch positions.

Draw an accurate sine wave - the gradient of the slope is not constant (or it would be a straight line instead of a curve) at the top and bottom of the curve the gradient is least and at the point where the wave crosses the x axis, it is at it's steepest. Converting the rotary motion of the POA s to a vertical motion gives changing rates of pitch change per unit of angular displacement.
If you still don't believe me then get your trigonometry tables out and look at the difference between the sines of 0, 30, 60 and 90 degrees - the change is NOT LINEAR and therfore not constant.
Helisphere - I must go to work but I wil try to answer you question Re Lynx tonight.

Lu Zuckerman

23rd Jan 2001, 20:41

To: Crab

We may be talking about two different things. One is flapping and the other is rate of pitch change. Using a pictorial analogy draw a circle. Divide the circle into four equal parts. The upper edge of the circle is the direction of flight. Where the right end of the horizontal centerline meets the outer edge of the circle mark that as point A=12-degrees. The opposite point on the circle is C=24-degrees. The top point is B=18-degrees and the bottom point is D= 18-degrees.

Addressing rate of pitch change; According to Sikorsky, if you broke the individual quadrants down to smaller elements which all meet at the center of the circle, and each of the smaller units were all equal, they say that the rate of pitch change from 12-degrees to 18-degrees is constant. And, if measured at the circumference of the circle the rate of change between elements of the circle would also be constant. The same would be true for points B-C, C-D and D-A.

Now draw three parallel lines spaced equally. The bottom left edge of the line is identified as 12-degrees and point A. Now draw a vertical line from that point to the edge of the top line. Now draw similar vertical lines spaced equally (2” apart). Draw a line from point A to the point where the second vertical line intersects the top horizontal line. Mark this as point C. Where the line intersects the middle line mark that as point B Now, draw a line from point C to the bottom of the second vertical line. Where the two lines meet mark this as point A. Where the line you made from point C to point A intersects the centerline mark that as point D. You can keep on going but the point I am trying to make is that the rate of pitch change is constant and that by graphing it out it can be seen that there is no sinusoidal curve. The line is straight.

Addressing rate of flap; First a couple of assumptions. You have a multi-blade rotor head that is fully articulated. The interlock number is much less than less than 1 so the rotorhead does not couple up with the disc path. Now, an analogy. When you shoot a bullet into the air it will go up and at some point that bullet will come to a complete stop just prior to reversing its’ course and start falling to the ground. Without scientific instruments it must be assumed that it did not fully stop and that the propulsive force just sort of faded away and gravity took over. I’ll try to apply this to the rotor and flapping.

Back to the drawing board. Draw a T. The top of the T should be 1-2” long. The vertical line should be the same length. At each end of the top of the T, make a large dot. That represents the offset hinges. From those dots draw a line about 3-4’ in length. These are your blades.
Now, draw the blades in a shallow V. These are your blades in a hover. Now, tilt the shallow V to the left. One blade will rise relative to the shallow V and the other will lower in relation to the shallow V. These are you blades when you are flying in any given direction. Using the circular reference above, mark the highest point on the shallow V as point D. The lowest point on the shallow V is marked point B. Connect the two points. This is your tip path. So, in flapping from point D to point B the blades are moving at a constant speed and discounting Coriolis forces and any drag forces the tip speed from point D to point B is also constant. The fastest rate of flap positional change would be between points D and A because when the blade reaches point A it is already changing pitch to a more positive angle. Regarding the bullet analogy the change of flap down to flap up although happening several hundred times a minute is gradual and not instantaneous.

Now, if "that" is what you said then I agree with you.

The only thing we don’t agree on is what caused the disc to tilt.

------------------
The Cat

[This message has been edited by Lu Zuckerman (edited 23 January 2001).]

JoePilot

24th Jan 2001, 00:07

Lu: your 'triangular sawtoothed wave' arises from the sinusiodal element being eliminated by having a horizontal axis which is non-linear (sinusiodally) and eliminates the sinusoidal nature of pitch change in your graphical representation.

Try plotting against phase! (instead of longditudinal position as you have done and wrongly interpreted the data)

Now all YOU have to understand is what makes the disc tilt...

(hope that saved you the trouble, Crab ... :) )

[This message has been edited by JoePilot (edited 23 January 2001).]

[email protected]

24th Jan 2001, 00:24

Thanks Joe - I am afraid Lu has lost the plot on this one - the Greeks sorted this business of Sine waves and circular motion out long agao but he doesn't believe it because it's not in the FAA manual he got for Christmas.

ShyTorque

24th Jan 2001, 00:54

I have got well behind the gist and I am having trouble following the discussion but:

Lu,

Are you saying that in the USA, helicopter rotor systems are controlled by applying forces to cause gyroscopic precession and that aerodynamic forces do not directly play a part in we in the UK call "flapping to equality"?

Lu Zuckerman

24th Jan 2001, 01:19

To: Joe Pilot and Crab

If you were to use a true sinusoidal wave it would mean that the closer you got to the peak of the wave either + or – the rate of change would decrease. And it would begin to speed up as you went on the downward side of the wave and then slow down as it approached the bottom end of the wave. That would mean that the rate of pitch change is not constant around the tip path. I was just graphing out what Sikorsky has been teaching for the last fifty years. They say that the rate of change + and – is constant. If you want to disagree with that please contact Sikorsky.

The fact that the description involved a saw toothed diagram instead of a true sinusoidal wave does not mean that the effects of the saw tooth wave will cause some sort of calamitous reaction in the tip path plane because, it doesn’t. The pitch change around the disc is seamless, as is the flapping around the disc. Assuming there are no mechanical defects in the blades or the rotor system these things happen and you are not aware of them because that’s the way the designers wanted it.

Now if you want to discuss the vibrations that result from flapping that is the subject of another thread. Most production helicopters have a means of combating that type of vibration.

------------------
The Cat

Lu Zuckerman

24th Jan 2001, 01:36

To: ShyTorque

Go to page 4 of this thread and several posts from the bottom is a post by Helisphere. Read that and then go to the top of this page and read my response to his post. Plain and simple, Flapping to equality is a term that is applied to Autogyros and it for some unknown reason has been carried over into the teaching of helicopter aerodynamics.

In the USA they teach that when the pilot pushes cyclic he creates an imbalance of lift across the disc. This imbalance is strongest over the left side of a counterclockwise rotor system. This perturbing force causes the disc to tip up over the tail and down over the nose. Aerodynamics initiated the force change but the actual movement was cause by the gyroscopic turning moment of the rotating disc.

If you look at retreating blade stall the cause and effect are the same. The left side of the disc is generating less lift than the right side. This imbalance will first cause the helicopter to roll left due to the disymmetry of lift and if you don’t catch it in time the disc will flap back or blow back pick one or the other. In either case you are dead or close to it.

------------------
The Cat

RW-1

24th Jan 2001, 02:04

>>This imbalance will first cause the helicopter to roll left due to the disymmetry of lift and if you don’t catch it in time the disc will flap back or blow back pick one or the other. In either case you are dead or close to it.<<

Well you got one out of three right.

First, you will have vibes, then usually you have a pitchup of the nose and a roll, either both at the same time or one before the other.

One can recover, as you have to be really ignorant of the developing vibes prior to the actual stall reactions in pitch and roll.

[This message has been edited by RW-1 (edited 23 January 2001).]

[email protected]

24th Jan 2001, 02:52

Lu, complete bo**ocks on two counts.
1. Having initiated a change in pitch cyclicly by pushing the cyclic that tilts the swash plate etc. the blades as they flap are seeking to recover the equal AoA they had before the cyclic pitch was applied. If cyclic pitch is removed then the AoA is reduced and the blade flaps down - in flapping down it increases the AOA back towards the original value. THIS IS FLAPPING TO EQUALITY - the equality is that of AoA. The disc assumes its new position with the tip path plane tilted forwards and stays in this new plane until something else (that is a gust or another control input and not the preaceesion fairy) changes the equality of lift(AOA) around the disc..
Moving into forward flight produces dissymmetry of lift because the advancing blade sees a change in rotational velocity which causes and increase in AoA - on the retreating side the opposite is happening - guess what... the advancing side flaps up with a high point at the front and the retreating blade flaps down with a low point at the back ----THIS IS FLAPBACK and is a result of flapping to equality. The pilot overcomes this with forward cyclic -something you can do in a helicopter but not in an autogyro - that is the difference.

2. Retreating Blade Stall - continuing from the above argument - as the forward speed is increased so the difference in AoA between advancing and retreating blades increases and to equalise the AoA the retreating blade has to flap down more and more as it's rotational speed is least. Eventually it reaches its stalling angle, has a rapid increase in drag and decrease in lift and flaps down. The effect is most pronounced due to the max rate of flapping between the 9:00 and 7:00 position so when lift is lost the aircraft pitches up and rolls left - THAT IS RETREATING BLADE STALL.

PS apparently one degree of cyclic pitch change produces about one degree of flapping.

Lu Zuckerman

24th Jan 2001, 03:58

To: Crab

This is another point where we must agree to disagree. You are 100% correct. But only on the eastern end of the Atlantic. If you came to the States to attend a class at Sikorsky you would have a very hard time to understand what the instructor was trying to teach because it would be totally alien to you. I have taught aerodynamics to US Army helicopter pilots and mechanics and I have taught in two A&P schools. They didn't have any problems understanding what I was trying to teach them. However if one of our guys came over to the UK to attend a course at Westland on the EH 101 and they taught principles of flight the way you are describing them they would have an equally hard time to comprehend the instructor because his words would be alien to him as well. Besides he would most likely be speaking in a midlands accent which doesn’t help matters much.

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The Cat

ShyTorque

24th Jan 2001, 04:36

Lu,

I have read same again, however I am still unsure if we are singing off the same hymnsheet or not. Please excuse me stating a few basics from a pilot's point of view.

Regarding the autogyro case, the disc flaps to equality as has been stated because it must be allowed to do so. As I understand, Cierva originally built small models from lightweight materials; they flew well. The blades were mounted directly onto the head, like sails on a windmill. When he went upscale to a full-size machine he used strong bracing wires above and below to control drooping and flapping of the blades due to their inherent mass / flexibility which he saw as a structural problem. He had unwittingly just designed a rigid rotor with no lateral cyclic control.

Unfortunately this was not the thing to do and feedback forces to the head and fuselage caused the aircraft to roll over during take off. He then realised that the small model's blades were so flexible they naturally flapped to equality by aerodynamic effects, without any hinging as such, which meant there was little or no rolling force fed back. He removed the bracing from the big aircraft and the fault was cured - it flew, with the characteristic back-tilted disc and the whole shooting match dragged along against its will by a big engine and propellor.

A helicopter disc will also flap to equality in exactly the same way if left to its own devices. Unlike the autogyro which has a separate forward thrust motor of some sort, the resulting flapback would cause the aircraft to go forwards for only a short time. The (now) rearwards horizontal rotor disc thrust component would now cause a reversal of aircraft direction. Obviously the pilot wishes to make forward progress and so makes cyclic inputs to make the disc FLY to where it needs to go to keep the horizontal rotor thrust component pointing forwards. With the cyclic held / trimmed forwards, forces are fed back through the swash plate to the fuselage resulting in a nose-down tilt to the aircraft.

Do you agree with this? If you don't then maybe there really is a fundamental difference of teaching depending on which school one attends.

BTW, I have no axe to grind here but having been taught and passsed on to many others the "British" way I am intrigued by all this argument over the very basics of how our machines fly. There will therefore be nil off-slagging from me!

JoePilot

24th Jan 2001, 18:29

Lu: Sinusoidal With Respect to PHASE

Lu:"If you were to use a true sinusoidal wave it would mean that the closer you got to the peak of the wave either + or – the rate of change would decrease. And it would begin to speed up as you went on the downward side of the wave and then slow down as it approached the bottom end of the wave. That would mean that the rate of pitch change is not constant around the tip path. "

That's right - now you are learning! ... I think if we came to an American school we would have no problem in understanding that the plot which you say the Sikorsky school uses is against LONGDITUDINAL POSITION - which you are just misinterpreting... simple really.

Crab and Shy express the other point well, try re-reading it with an open mind - then you'll learn more...

(There's just a slightly subtle error with crab ... the AoA's aren't actually equalised ... because the airspeeds are different Clv2 same thing really tho')

Helicopters still fly in the US because none of this really matters....

offshoreigor

24th Jan 2001, 21:39

Lu:

I once said on a previous thread:

"If your talking, You're not listening. If you're not listening, you're not learning."

Lu, you talk too much!.

Cheers, OffshoreIgor http://www.pprune.org/ubb/NonCGI/eek.gif

[email protected]

24th Jan 2001, 23:14

Lu, as ever, you have misinterpreted a diagram designed to highlight the fact that pitch changes throughout the 360 degree orbit and guessed that they really meant to show you that pitch change was constant. If that was what Sikorsky had meant to show, they would probably have used a 44 sided shape and not a circle. The 44 sided shape with pitch change constant would produce your saw-tooth wave when drawn - the circle produces a sine wave which for the second time does not have a constant gradient and is therefore not linear and does not give constant rate of pitch change.
As a further counter to your precession theory - you happily accept that phase lag changes with hinge offset - Lynx and BO105 have phase lag of between 75 and 80 degrees - now show me a gyro that precesses at anything less than 90 degrees!!! Oh no you can't.
Furthermore I have just been looking at a book that is imaginatively titled FM 1-203; it is produced by the US Army and is the fundamentals of flight manual issued at Ft Rucker. It has pages and pages of flapping to equality, non-uniform rates of pitch change around a control orbit, blowback etc etc etc and one 5 line paragraph about precession. So much for vive la difference across the Atlantic then. Are you sure the book you read was written after the Wright Brothers got airborne at Kittyhawk?

PS try explaining why the coriolis force effect that deflect air moving in the Northern Hemisphere to the right is negligible at the Equator and increases as you increase Latitude......hmmm, might it have something to do with Sine of the Latitude changing non linearly?

rotorque

25th Jan 2001, 16:08

I have been following the last few entries with a little eye brow raising. The one thing that I have seen that gives concern is the use of the terms 'pitch change RATE' and 'pitch change ANGLE'. One or other of the two will infact be constant while the other will change throughout each revolution.

Are you sure that you are not getting the two mixed up? You may be arguing the same point.

[email protected]

26th Jan 2001, 00:17

Lu,

Do you think the piston in a car engine moves from top dead centre to bottom dead centre at a constant rate? Of course not - the rapid burning of the fuel expands the air and ACCELERATES the piston downwards - the piston stops moving vertically at the top and bottom of each stroke - WITHOUT tearing the engine apart.

To use your hydraulic pump analogy - the fuel pump in the Wessex was exactly the same with pistons moving up and down following the swash plate; the angle of the swash plate determined the stroke of the piston and thus the output. My point here is that the rate of stroke of the piston is not constant - it is accelerated downwards as a very efficient method of compressing the fluid - fuel in this case. As a piston on one side of the swash plate is accelerating downwards, the corresponding piston 180 degrees out is accelerating upwards so the net effect is balanced and the pump does not tear itself apart. The piston begins to move downwards at the beginning of the stroke, reaches its maximum rate of movement mid stroke and then reduces its rate until at its low point where the reverse process begins.

Back to helicopters - the Lynx disc is low at the front when you push the cyclic forward - the advance angle is 15 degrees - the pitch actuator is mounted 15 degrees ahead of the lateral line through the rotor mast - this is because the phase lag is approximately 75 degrees (15+75 =90) ie it only takes 75 degrees of angular rotation from the application of the force to achieve its high or low point - CAN A GYRO DO THIS?

The Lynx head does not have a mechanical flapping hinge - the titanium forging flexes to allow the flapping to occur and gives an effective hinge offset of between 12 and 17% depending on who's book you read. The fact that phase lag reduces as hinge offset increases is well documented - WHY this happens I do not fully understand as I am allergic to greek flute music and have to take a day off if I see more than one variable in a formula. I am sure you will have a lengthy explanation.

ShyTorque

26th Jan 2001, 09:38

Lu,

I am intrigued by your answer to my last post. Presumably we are merely using the same language in a different way?

How can a change in airflow over a rotor blade cause a gyroscopic force? In my understanding airflow over an aerofoil causes lift and drag, not gyroscopic forces.

I certainly agree that there are very large gyroscopic forces involved in a typical rotor system but occurring as a result of masses being rotated at speed and not as a result of blade pitch angle changes. Various (enormous!)figures have already been quoted.

My understanding (as a pilot with 22 years of rotary experience and much of that spent instructing) and that of the conventional British way of teaching is that the rotor disc is "steered" not "precessed" by using aerodynamic forces to overcome gyroscopic forces already acting. Aerodynamic and gyroscopic forces come to a compromise and what is left over gets fed back into the rotor head. If the resultant feedback of the two is too great for the pilot to handle then it is usual to design a hydraulic system to help him out.

My logic tells me that the relatively spindly pitch-change rods on a typical rotor head would be totally inadequate to apply sufficient precessional forces at an acceptable rate for purely gyroscopic control of a rotor disc, as would the pilot's own physical strength or the small hydraulic pump as typically used.

Maybe I am missing your drift due to us being from two different nations separated by a common language!

[email protected]

27th Jan 2001, 00:33

As ever Lu you have failed to follow anyone elses logic but your own.

The reciprocating piston engine converts linear motion to rotary motion - the pistons move up and down accelerating then stopping because that is what they must do do rotate the crankshaft at a constant angular rate.
Even a schoolchild could understand the reverse of this process when a circular object rotating at constant angular rate converts this motion to vertical movement by accelerating then stopping a piston or a Pitch Operating Arm (back to the original point I think)

Yes the Lynx rolls to the left following the application of forward cyclic but only as a result of an acceleration cross couple and not the simple rigging error you attribute to the R22. The reason the Lynx fuselage reacts so quickly is due to the effective flapping hinge offset that gives it huge control power and is the secret of it's manoeuvrability (sorry, too many vowels in there for American readers).

The two helicopters you describe are different sizes but the important feature you fail to grasp is that of hinge offset. On a teetering head the flapping is allowed acroos the pivot of the rotor mast - on an articulated head the designers can put the flapping dragging and feathering hinges where they like. The hinge offset is where the flapping hinge is physically located expressed as a pecentage of the distance between the axis of rotation (rotor mast) and the blade tip. I suspect that despite the disparity in size, both your Sikorsky's have equivalent flapping hinge offsets that are probably quite small - therefore phase lag is very close to 90 degrees and the 45 + 45 combination of pitch horn lead and swashplate offset works well.

As hinge offset is increased the blades take less time to reach their high point approx 75 degrees in the case of the Lynx which has an effective (not physical because the flapping occurs in the titanium forging rather than an actual hinge) hinge offset of between 12 and 17% - the figure vary from book to book as there is more than one way to calculate effective hinge offset.

You could have left all that rubbish about rotary engines out as it proved exactly diddlysquat!

[email protected]

28th Jan 2001, 23:42

Joe Pilot,

Just reread your post at top of this page - sorry I should have mentioned I was trying to keep things simple for Lu and only consider the effects of cyclic flapping without going into forward flight as the change in V2 would have complicated my argument. Very glad to see someone actually pays attention to my drivel!

RW-1

29th Jan 2001, 02:27

>>(sorry, too many vowels in there for American readers).<<

I cannot help but feel hurt http://www.pprune.org/ubb/NonCGI/frown.gif

Not all of us fall into that basket .... Some of us keep a dictionary in the loo just for reading :)

[This message has been edited by RW-1 (edited 28 January 2001).]

[email protected]

30th Jan 2001, 23:35

RW-1, not an insult, more an attempt at humour(humor) to highlight our separation by language, viz. Manoeuvre v manouver and number of vowels.

RW-1

31st Jan 2001, 00:17

Darn Crab, that was my attempt too ... time for us to tune to the same freq :)

By now I though you would know when I'm hurling hehe ....

Lu Zuckerman

31st Jan 2001, 01:55

To: RW 1
I noticed that you pulled the plug on your 407 thread. Were you afraid if someone asked for my theory on why the tailrotor hit the fuselage and/or a theory on what would cause a blade to diverge from the plane of rotation and hit the fuselage that it would hijack you thread?

Regarding your comment about taking a dictionary to the loo it is obvious that you don’t avail yourself of its’ contents based on your spelling in your posts. Could it be that you use its' pages in place of toilet tissue. That was an attempt at humor/humour.

HeHeHe
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The Cat

[This message has been edited by Lu Zuckerman (edited 30 January 2001).]

RW-1

31st Jan 2001, 04:06

Lu Zuckerman: There should be a NOTAM for you, pilot wannabee.

I pulled the posts for reasons that are much too complex for your grey matter (based upon your history) to comprehend.

Your statement is yet another 6 year old goad, perhaps it will be pulled as well as the other crap you tried in another thread?
It is any wonder to you yet why my response stayed, but yours gets pulled?

Your going down your hole even further...
Dante himself now looks down upon you ...

And it is no suprise, your comments, I sent several members a likely response you would give when you found the thread gone, so far? I was right on.

No Lu, I pulled it because YOU failed to JUST give your theory on the tail strikes.

You wanted to banter about on whether one mishap was caused by it or the mains, instead of just addressing the AD, which was the purpose of MY thread.

A classic Hoverman example of you just posting to:

Hear yourself talk.
And start something with others.

There was no reason for it, and you have no excuse to account for it, again, your history preceeds you, and the others have already formed their opinions.

But as evident in your history here, you never just get to it, you strive to create ill will, illicit responses that you will no longer get from me.

Rather than subjecting this whole forum to yet another example of your driveling, I did everyone a favor and removed it.
If medals were given for letting others avoid you, I'd get the Congressional medal of honor for that move.

Your humor is as weak as your intellect, and you have already shown everyone the truth of that statement; all anyone has to do is read your priors.

I don't have to attack you any more LU, you do a better job of foiling yourself than I could ever dream of doing. I'm sure that irks you to no end, a lowly PPL seemingly getting better responses than the houty touty multi-degree engineer.

Truth is it's getting (or really gone past) the point where the real posters here just don't care what you have to say any more.
And the minority (would have been the majority 6 weeks ago, but not anymore)is getting smaller everyday.

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Marc

Lu Zuckerman

31st Jan 2001, 05:41

To: RW 1

“No Lu, I pulled it because YOU failed to JUST give your theory on the tail strikes”

Typical of your response to everything you obviously did not read what I stated in my post. I responded to your request for my theory on the tailboom strike and the possible main rotor strike. I ended my statement with, “any takers”. I expected to get a positive or a negative response. Some people may have been interested because of a vested interest in the 407. However before anybody could respond, you had pulled the thread.

“It is any wonder to you yet why my response stayed, but yours gets pulled”?

None of my responses were pulled. I removed them hoping to defuse the attacks on myself but nobody seemed to notice and the threads without my posts took on a life of their own where every body kept slagging me off as a person.

Now, I will ask you in the presence of all of the onlookers, do you want me to state my views and/or theories on what might have caused the tail boom strikes and the main rotor strike on the last 407 to crash? If you do, open a new thread, as I don’t want to be constantly reminded in my email that someone has responded on my thread. You can introduce the thread in your own inimitable fashion. You can even call it Lus’ Vues on the 407. The ball as they say, is in your court.

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The Cat

RW-1

31st Jan 2001, 06:26

>>. I responded to your request for my theory on the tailboom strike and the possible main rotor strike. <<

No, you did not. You offered to do so after 5-6 round and round semantic posts as to whether the latest mishap was related to the AD and whether the main rotor strike was the cause or secondary.

Are you so that wrapped up in yourself to believe that? My my, you truly are a piece of work.

Again, don't confuse yourself in your ongoing quest to hang yourself.

>>I removed them hoping to defuse the attacks on myself but nobody seemed to notice <<

And the reason is for the above, my posts on the top of page 4, 5 and above.

You figured out just how much that post weakened your already precarious position on the board, one step in the right direction, but don't expect anyone to believe you, as stated before and not by me, your credibility has lost a great deal.

There is no ball, if you had acted professionally to begin with, I wouldn't have had to pull the thread. Another 6 year old dare.

>>Typical of your response to everything you obviously did not read what I stated in my post. <<

This is yet anotheer LU' ism that hoverman failed to list in his imatation of you, you like to use that one way too much.

I believe you should use that on yourself anbd ask why you would continue to post here at all, ask why so many of your own supporters 6 weeeks ago are now as fed up with your posts as I?

Nothing more needs to be said, you don't get it, and never will unless someone put's it in a textbook for you.

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Marc