I once made a homemade Gmetre and got up to 17 G's. I know more is possible but what's the maximum yank possible on our models? 17 is a pretty good yank and I now design my aerobatic models for 20-25 G's and this has worked well in practice. I thought I remember someone talking about calculating it but I'm short a few cells in my thinker to pull off that kind of stuff!

Hi Tapio,
what size plane was it that you got to 17 g's? I'm curious as to how much the 35% planes do. Also how did you make your g-meter? - sounds cool
later,
Jon

R/C model sailplanes doing dynamic soaring have hit a measured 170 MPH while pulling tight circles. It is estimated that they have pulled about 40 G's in the process. Several have exploded into composite confetti in the attempt.

See:
http://ourworld.compuserve.com/homep...tone/dsoar.htm

Ollie...
Thanks for posting that DS web site. The explanation and illustration are just great and, very interesting.

Bill

Yes I've seen a DS soarer blow up on the radiocarbonart videos.
My friend flies a Diamond dust with a hot 40/pipe and pulls hard
coming straight down full throttle. He often lands with no covering
left on the bottom!!!
Rebel the metre a made was crude. I tested some alum. tubing to
the bending point to get an oz/in figure.Then as an example one
was to see if 20 g's was exceeded simply put one twentith weight
on a supported tube in the plane.
A real seaplane builder said a real g-metre is small and simple.
The plane I pulled 17 with was a 2-metre Falcon glider,RG-15 wing
with a good spar vertical dive with a 25 on a pipe. Probably pulled her out at 100 mph+-
On 30% planes the speed is lower but I'm guessing I've pulled close to 15 with my 80" Extra. I do snaps at full bore on the level
and a blender with power on,once blew the top of my rudder off!!
Anybody know how the tailfeathers are loaded in flight???

The nose down pitching moment of the wing increases as the square of the airspeed and is closely related to the camber of the wing airfoil. The down load on the stab that balances the pitching moment of the wing is inversely proportinal to the tail moment arm length. The pull up in the zoom just after release could produce a download on the order of 10 pounds. For slope racers with elevator to flap coupling, the down load on the stab in a tight turn could be even higher. For DSers pushing 200 MPH the down load could easily be in the range of 20 to 30 pounds.

It is easy to figure the G loading in a coordinated LEVEL TURN, if you have a calculator with trig functions.

Just use this formula:

1
------- = G , where G= g loading, and B= bank angle
cos B

So, on your calculator, punch in the bank angle, let's say 60 degrees. Next, hit the "cos" key. Then, hit the key marked "1/x". The number now in the screen is your G loading, which in this example, should be exactly 2.

Keep in mind this is just for a level turn (not climbing, not descending)

If you try to calculate G for a 90° bank angle, you will get an error, because at 90°, G could theoretically go to infinity. Also, a coordinated level turn is impossible at 90° bank angle.

G does not vary with airspeed or groundspeed for that matter either. For example, a 60° level turn is 2g. It is the same whether your speed is 30 mph or 200 mph. Your turn radius will increase with speed, but G will be the same.

I just thought this would be an interesting bit of knowledge to pass on to the non-engineers out there...(like me... I learned this in Private Pilot Ground school).

The most basic things that determine how many G's a plane can pull are:
1. The strength of the airframe (particularly the wing).
2. The flutter resistance of the airframe.
3. How fast the plane is going.
4. The maximum lift coefficient of the wing.
The number of G's is just the ratio of the lift force to the weight.
The lift force is proportional to the wing area, the coefficient of lift and the square of the airspeed.

A special case is a flying wing in which the mass (load) distribution matches the lift distribution. Such a wing has no bending forces, no matter how many G's are bening pulled. It will not break due to spanwise bending.

Ollie - when you talk about g-loading and max coefficient of lift, max g-loading would be the ratio of max Cl / operating Cl (If Clmax were 1.4 and operating Cl is .4, g-loading would be 3.5)When pulling quickly, however, that theoretical max loading can exceed due to the momentum of the air and something called "hysteresis. Are you familiar with that and is that why some aircraft are pulling so many g's?

Yes, if the speed doesn't decrease in a level turn, the maximum G loading is the maximum Cl of the wing divided by the Cl in level flight at maximum speed. For a pylon racer the Cl in level flight is usually much less than 0.1 and the induced drag coefficient is usually much less than 0.0001.

The Hysterisis I'm familiar with occurs after the wing stalls.

I had been told by a professor once that if the pull was very rapid, the momentum of the air would allow attached flow at an angle of attack greater than the stall angle of attack with (for a brief period of time) and, hence, a higher-than-normal Cl. That is one reason it's a very bad idea to check if Va actually works with an abrupt pull of the stick (yoke) as Va is a theoretical value determined with a load factor of (normal category) 3.8 g's and a "steady-state" maximum Cl.

The S6063 airfoil has a nice stall margin above the top Cl in the low drag bucket.
The top of the low drag bucket is a Cl of about 0.4 but the airfoil doesn't stall until Cl of 0.8. You can calculate the bank angle for Cl 0.4 at speed and have a fairly reliable visual indication to the pilot of the G's to pull while staying within the low drag bucket. This behavior is typical of thin airfoils. This avoids the transient and other detrimental effects of pulling too high a coefficient of lift.

Ollie,

Your comments in orevious posts re the 6063 got me interested in giving it a look see. I loaded the coordinates for both the 6063 and the MH33 that I have used quite a bit. I have a pretty good airfoil program that will let you masage the foil then you can print templates etc. It's a freebe that I picked up a while back. So, I sized both the mh33 and the s6063 to the same cord and printed them out for a comparison. Big shock to me! These foils are so close that they are very close to being identical.

Next I loaded them into CompuFoil. CF reports that for the 6063 the zero lift angle of attack is 1.36 deg. For the mh33 it's 1.22 deg. Verry close.... Ideal angle of attack for the 6063 is 0.37 deg. for the mh33 it's 0.39 deg. So, would it be safe to assume that if the engine and stab are at zero zero that the wing should match the "ideal" as stated by CF? What's your take on this "ideal angle of attack" reported by CF? Rgds, Jeff.....

The angle of attack is not determined by the incidence relative to an arbitrary line drawn on the plans. Incidence is easy to measure. Direct measurement of angle of attack is best done in a wind tunnel. Direct measurement of the angle of attack of the wing of a model in flight has rarely been done and requires special instrumentation.

A plane in level flight at a constant air speed takes on a pitch attitude which produces the necessary angle of attack. In this state of dynamic equilibrium, the sum of all the moments is zero and the sum of all the vector forces is zero. A simple case would be where the lift acts at the CG, the down load on the tail times its moment arm is equal and opposite to the pitching moment of the wing, the thrust line passes through the CG, the lift is equal to the weight and the thrust is equal to the drag. For the lift to be equal to the weight at a certain airspeed, the wing must be at a certain angle of attack and the plane assumes a pitch attitude by the flight trimming process that produces the necessary angle of attack relative to the direction of flight relative to the air mass it is flying in. If the plane changes airspeed and assumes a new state of dynamic equilibrium, it will assume a new pitch attitude to produce a new angle of attack such that the lift produced is equal to the weight. The new value of thrust will produce a new airspeed such that the new drag is the same magnitude as the new thrust. The new pitching moment of the wing matches the new down load on the tail.

Therefore, the "ideal angle of attack" is determined by the throttle position and the elevator trim position for a given CG. The plane then takes on a pitch attitude which produces the "ideal angle of attack." The trick is to arrive at the airspeed, CG location and elevator trim that produces the "ideal angle of attack" for that model. The "ideal CG location" will result in zero down load on the tail at the associated airspeed.

When flying in a level turn the trick is to fly at the angle of bank that puts the wing at the "ideal angle of attack" for that airspeed and lift force. The desired angle of bank can be determined by flying circular laps at full throttle and various angles of bank to see which angle of bank produces the fastest lap time.

A paper or computer analysis of a particular design configuration can give you an idea of how it will perform and the configuration can be changed to see if the performance improves. In this way the designer can evolve a design that is more refined. The main problem is that the coefficients of parasitic drag for exposed cylinder heads, control horns, etc. are not generally known because they haven't been measured.

The engineers at the University of Stutgart did some measurements and found that the exposed control linkages for two flaps and two ailerons on an F3B model were about 10% of the total drag in high speed flight!!! The message is that attention to detail in streamlining can have very benefical results.

BTW, I chose the S6063 in my example because it seemed to fit the application and the polars from windtunnel measurements were available at the reynolds number of interest. The particular wind tunnel test model was not an accurate representation of the airfoil because it had a slight reflex in the trailing edge of the model. The mean line camber, lift and drag coefficients were a bit lower than the true S6063. Since Martin Heperle's web site is shut down, the MH33 polars generated by X-foil were not available to me.

There where inflight test made in germany with a 3D-piezo accelaration sensor and a datalogger in a 95" aerobatic plane with G62 or ZDZ80. Normal smooth maneuvers showed the expected Gs of 2 to 6 Gs but the aggressive positve and negative snappy maneuvers showed Gs up to 30!
The reason for those testflights where some broken wing tubes which already caused lawsuits.

I read some where that r/c planes rarely pull more than 10 g's but i may be wrong. I will have to remember that formula to.