Static Load Testing

Discussion in 'Aircraft Design / Aerodynamics / New Technology' started by Jman, Feb 4, 2003.

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  1. Feb 4, 2003 #1

    Jman

    Jman

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    I figured I would start this thread of discussion to get some dialog going about static testing aircraft structures. I've been following a thread over in the cutting edge discussion about carbon spars. You can see the thread here. It's got me curious as to the best method for static testing spars and wings. Does anyone know of a good resource with practical tips for setting up a rig to do this? I did some searching around the net but did not come up with very much. I figured that maybe some of you might have some insights on the matter. In the other thread, Senna posted a diagram of his proposed testing method and to me it looks great. I just figured that the testing methods deserved it's own thread and would be an interesting topic to explore. Hope you don't mind Senna :D.

    Jake
     
  2. Feb 6, 2003 #2

    Senna

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    Heck no... I don't mind a bit. Having a forum where we have the ability to exchange information, benign points of view and even a vociferous opinion or two is a tremendous resource. It's a bit slow here at the present and until we get sufficient mass (aka members) to experience the flywheel effect... we NEED topics of discussion.

    I visited Jim Maupin once down in Southern California many years ago. He was working on the WINDROSE at the time. We got on the subject of the uniqueness of the wing spar design and apparently Jim and Irv Culver (Irv designed the wing) were of differing opinions as to whether the lack of a true shear web between the caps would be strong enough. Well... they static loaded the wing (which I got to witness) and although it didn't fail... the test showed they needed something more than just the foam between the caps. They ended up placing vertical wooden dowels down into the foam at intervals so once the caps were epoxied in place the dowels would tie the upper and lower caps together.

    So this is what I was able to gleen from the discussions I had that day with Jim and from the few sources I’ve managed to find about static load testing... in no way is this intended to be a definitve explanation.

    Apparently you need three things for static load testing. The spanwise load distribution for the particular wing planform, the test fixture itself, and sufficient weight necessary to complete the test. The test I witnessed used 50lb bags of PLAY sand... the stuff you'd put in a child's sandbox.

    You can’t just pile weight on a wing and call it a sucess if it doesn’t break. The load needs to be distributed across the surface of the wing in a way that accurately reflects the loads the wing will experience in flight. It’s important to calculate the load distribution that accurately approximates the flight conditions you’re testing for. I’ve heard that some test the wing at a high angle of attack and load it to simulate the Va min or the minimum design maneuvering speed. I don't fully understand this particular test but apparently at Va min if you do a max pull-up or max bank angle maneuver the wing experiences the loading at a higher angle of attack. So to proof to the wing it is mounted at the high angle of attack to simulate statically the same loads it would in flight. However most static tests proof the wing at a lower angle of attack for a maximum load factor such as those induced at Vne, the speed to never exceed.

    The actual test is fairly simple. The wing gets bolted upside down to a support structure and the loading begins. It’s important the test stand be built in a way that it does not add any extra strength to the wing you’re testing. While the bags are being loaded, the wing tip is supported. When the desired load is reached, the supports are removed and the wing is now supporting the load, reacting as it would to a flight load. The amount of structural deflection is measured, and photographs are taken for documentation. I’ve seen placards included in the photographs having the date and amount of weight. The wing must support the load without failure or no permanent deformation for 3 seconds. The supports get re-inserted, and you either go to the next higher test weight or the weight is removed. The wing structure should get a critical inspection for anything that may have gone unnoticed.

    Like I said this is more an overview rather than definitive. I've no formal schooling to support any of the above... but at least it's a start for a discussion.

    Senna
     
    Last edited: Feb 6, 2003
  3. Mar 3, 2003 #3

    orion

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    Your overview is concise and hits on the right points. Static testing can be simple however it is rather important to examine all the anticipated critical points. Examining the wing for different angles of attack is not always critical but it can show useful behavior characteristics of the wing under load.

    High g loading of an airplane's structure usually occurs at high angles of attack (wind flow relative to the wing's chord plane). Testing the wing in a horizontal position examines the spar in simple bending but at a high aoa more of the wing's structure is at play including the skin and the aft spar.

    It's also a good idea to be as conservative as possible when running these tests. Most folks calculate the loading based on an ideal elliptical distribution. This is OK however it is strictly applicable to only a few actual planform layouts.

    It is better to use a mocified elliptical distribution, with a slight emphasis of the load out toward the tips. This has two benefits. first of all, if the wing passes then you'll know the structure will be applicable to most phases of flight, regardless of the actual load distribution.

    The other benefit is that everyone out there has a slightly different building technique. Some are better than others. If you design a conservative wing that can pass a pessimistic load distribution, then you have a good chance of delivering a wing that is sound even for someone that may not build it exactly perfect.

    The weight penalty of this type of structural approach is very slight however the peace of mind is substantial.

    If you're just building for yourself, extra structural ability is again more peace of mind.

    Finally, it is also important to not limit this discussion to just the wing. Remember, to reach this loading on the wing the rest of the airplane is affected also. This covers the load on the firewall by the engine and prop (don't forget the gyroscopic load for the latter), the loads on the interior structure by the occupants and the wing attachment, and of course the loads of the tail surfaces and their effect on the fuselage.

    Testing the wing is only a part of what needs to be examined.
     
  4. Mar 19, 2003 #4

    dannicoson

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    There are several load points that should be considered when designing and testing. These are by no means the only ones.

    1) Maneuvering Speed, Va: By deffinition this is the speed at which you can make full throw control inputs and not overstress the aircraft. What this really means is that you are stalling the main wing just as the airframe achieves max allowed G. This also means that the wing is at the maximum angle of attack, usually about 12-15 degrees for most airfoils. I don't think it would be worth trying to mount a full size wing at that angle, I think you would risk damaging the wing from awkward mounting and loading in that position. And actually if you think about it, the actual aerodynamic load is still normal to the chordline of the airfoil, Lift is defined as the vector portion of this force that is normal to the freestream flow. Just math, so don't worry about re-creating the high lift AOA.

    2) VNE velocity never exceed: It's been a year since I researched this so... main point here is that as you airspeed goes up (coefficient of lift goes down) and the center of pressure (Cp) moves aft. This is true for any cambered airfoil which most of our homebuilts use. This means for a high speed aircraft like a Glassair, Lancair or even a C-172, at the top end of allowed airspeed you can reach max allowed G (load) at a very low angle of attack and the center of pressure is well towards the center of the airfoil chord line. What this means is that as much as half the load will be carried by the rear spar on a two-spar wing. So in this case you need to verify that the rear spar is actually carrying it's share of the load. On my own design the rear spar carries half the load in this condition. So your static load test would need to ensure the wing is loaded to simulate the actual air loads...

    3) Max Flap speed: What is your max G load at max flap speed? With big Fowler flaps hanging out in the wind you are again loading up the rear spar very significantly. Can you simulate this load?

    As mentioned in some of the earlier posts many other parts of the aircraft need serious attention. I dug into FAR 23 last year and found it to be a treasure-trove of design guidance when it came to sizing loads and allowing for gust loads etc. It IS like reading tax code but it is very comprehensive and following those guidlines makes me feel much better about any design decisions.

    Dan
     
  5. Mar 19, 2003 #5

    Jman

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  6. Mar 19, 2003 #6

    orion

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    Good points in the previous post however I need to add a bit of correction. While airfoil information does show a max angle of attack prior to stall of 12 to about 16 degrees, that is applicable only to the two dimensional case. A three dimensional wing develops stall far beyond that point. Actual stall of a three dimensional wing can reach almost thirty degrees although twenty to twenty five degrees is more common.

    I also don't think that recreating a high aoa for testing is all that critical however, if you are looking for the behavior of the structure at the extremes, you may have to. It is important to keep in mind when designing the structure that at those angles of attack the pressure distribution shifts far forward, creating a different loading than seen in level flight - this condition can be approximated by the load distribution of your test weights in level attitude so again, the high aoa set-up may not be necessary.

    As far as the Cp shift is concerned, for most general aviation applications the magnitude of that shift is relatively small and so for flight purposes, can be easily trimmed out.

    For conventional structures, a simple configuration places the main spar in bending and the skin in torsion. This combination can easily account for the variation of the Cp position. Placing the aft spar in bending is also a solution but unless it is designed correctly, it can direct the stress distribution to a part of the wing you may not want loaded.

    The classic way to design a wing is to pin the aft spar at the fuselage (placing it in shear only) rather than attaching it so that it is in bending. That way all the aft spar has to do is maintain the shear integrity of the skin attachment, and of course carry the control surface and flap loads. Under high loading, the aft portion of the wing is allowd to deflect. Since the deflection is in the form of washout (under positive load), the resutant twisting motion of the wing actually unloads the tip thus reducing the bending moment at the root and also the chance of tip stall.

    As far as the Cp shift is concerned, yes, it does move aft however it does not start reaching the 50% chord point until you reach high subsonic Mach numbers. Remember the P-38 tuck-under problem? That was an issue of Cp shift. When the aircraft accelerated in a dive, the Cp moved so far aft the controls became difficult to move. Beyond a certain point it was virtually impossible to pull out of the dive. As the aircraft accelerated more, the Cp moved even further aft creating a high nose-down pitching moment, tucking the nose of the airplane into an even steeper dive. The eventual fix was dive brakes but not before several pilots lost their lives.

    The speeds involved however are very high and so, for most general aviation applications, this is not so much of an issue. I personally have seen a slight stiffeneing of the controls in a G-III in a dive, but this was well past 300 mph and the condition was still very controllable.

    One last point - as the last post indicated, Part 23 is an excellent reference especially if you have the time and can keep your eyes open long enough. One of the best sections for general aviation though is Appendix A, which is a simplification of several loading issues for aircraft weighing less than 6,000 pounds. It does refer to the main body for several items such as gust loading, but the simplified presentation is much easier to understand and does not require as much digging.

    Another good source for reading is the European JAR-VLA document. Basically it is similar to Part 23 but it is written particularly for light airplanes so a lot of the extra unneeded infomration is left out.
     
  7. Mar 21, 2003 #7

    Holden

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    Orion,

    The Mach tuck issue had not only to do with the CP change, but also the tail surface stalling due to the shock wave. The flow would separate from the tail and the tail trim force would decrease.

    The real "fix" to the problem was the all flying horizontal. This was kept "top secret" for 15 years after Chuck Yeager and the Edwards folks figured it out. I had a professor who was the staff Phd at Edwards and in the tower when Yeager "broke" the sound barrier. Switching to the flying horizontal "fixed" the problem, NOT adding speed breaks. The P-38 fix was not a fix, just a way of getting out of a bad situation.

    Mach tuck is an issue on airliners and corporate jet today. Most have a horizontal that is "trim" fixed with flap trailing "elevator." A Kitfox has the same setup. The way to pull out of Mach tuck is to "trim" the leading edge down fast. The "elevator" is in the wake of the shock-stall and the more it is deflected the more it stalls and looses lift. The only way to trim is to rotate the slab.

    This "Mach tuck" situtation also can occurs during ice up of the horizontal and/or main wings. There is no shock wave, just a stalled horizontal and main wing in low angle of attack. The main wing can stall from ice and the flow separate causing the need for nose up trim fast. The "elevator" causes the flow to separate more as the elevator is deflected.

    The "fix" for ice is also the all flying stab, or a leading edge trimable surface. The problem with the leading edge is that the response time is too slow, and it requires training. A pilot must be trained to trim instead of pull back when the nose drops.

    In both the Mach and ice cases, the main is also in a stall, and lift is lost, usually 25% or more. The airplane drops until that lift is regained, adding to the need for a quick response.

    So, I use "all flying" horizontals with an anti-servo in my design for ice safety reasons. My hull is also placed to counter the pitch change when the main wing stalls due to ice. The forward gear door can also aid in pitch trim if needed. The rudder placement/coupling allows for roll control.

    Perhaps one of the most important issues in making small airplanes useful to the general public is this ability to stay in control when the wings ice up. As you know, de-ice equipment requires training and can fail. Having the airplane naturally ice resistant is key to any wide acceptance of the vehicle as transportation for the layman. The solution is also the same solution for the "problem" encountered in Mach tuck, in part.

    Holden
     
  8. Mar 21, 2003 #8

    orion

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    Hi Holden;

    Thanks for the rest of the P-38 story. My information came from a couple of folks who I worked with at General Dynamics about eighteen years ago. One of them used to be an aerodynamicist for Lockheed through much of the war period. He indicated that the tuck-under problem was primarily a high subsonic Mach Cp shift, which resulted in overly heavy control surfaces. I do remember him saying something about the Mach stall, although I do not recall the details. It however does make sense.

    The part of the story I got was that several pilots manged to get out of the dive by extending their gear. This apparently slowed down the airplane sufficiently to recover the control authority. Based on this information, the manufacturers for a time added dive brakes to the airframes.

    When did the flying stabs get added? I don't remember seeing any with anything else except the conventional horizontals. The last two I saw were the "J" and "L" models and while I don't recall the horizontal configuration of the "L", I'm sure the "J" had the standard stab/elevator layout.

    Bill
     
  9. Mar 22, 2003 #9

    Holden

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    Orion,

    You are correct that the "fix" was dive breaks for the P-38. What I said or tried to say was that the real fix was, or should have been, an all flying stab. Engineers did not know this until after WWII when the Edwards guys figured it out. My Professor, along with a "Discovery Wings" channel program I saw a year or so ago stated the same idea and story line, although they did not mention the engineers behind the scene.

    From what I understand, emergency procedures call for extending the gear also, or basically anything to slow it down, in a do or die fashion.

    If you search for "mach tuck" on Google there are several articles, one of which mentioned the P-38.

    http://www.flightjournal.com/articles/me163/me163_6.asp
    http://www.ultralightfloats.com/machtuck_article.htm

    My interest was in understanding how this can help, if anything, in making a <.35 Mach airplane fly safer during ice up, and during a normal stall. I searched the NTSB for ice up crash reports and found a common theme of a dive in and tail loss that would appear to fit the same Mach tuck effect, except the cause was ice. That is my theory, and I have not seen anybody else make the connection. In the Mach tuck, the shock wave causes the flow to separate, and with ice, the ridge causes it. Effect is the same...two stalled wings.

    For an airplane to hold pitch with two stalled wings requires an aerodynamics similar to many of the new fighters with forward fuselage side "wings" that counters the strong negative pitching caused by the stall and that maintain total lift as the wing stalls.

    In other words, what every pilot is taught about the stall must be changed by design. The typical stong pitch down found on a GA airplane must be eliminated. The airplane must hold lift and have only a gentle negative pitch and be able to maintain lateral control in a post stall. If the airplane is designed as such, the airplane is intuitive to the layman, and will maintain lift in the stall as AOA increases past the stall point. The result will be far fewer accidents and far less training required, along with far fewer ice related crashes.

    A mono wing or a canard cannot accomplish this. In order for the layman to fly, engineers must move beyond the standard mono wing or canard and look with the right goals in mind. The Mach tuck issue just happens to be a major clue that engineers have overlooked for GA.

    The new generation of personal jets that will be falling out of the sky with layman pilots, are not able to deal with Mach tuck, ice or the stall. If the engineer solves the Mach tuck he can also solve the ice and stall issues, along with the training issue. All four are related. No new personal jet I see deals with this key issue.

    Holden
     
  10. Mar 22, 2003 #10

    dannicoson

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    Orion,

    In my earlier example of Cp shifting aft as airspeed climbs, I was stating that 50% of the load in that high speed condition was being placed on the rear spar of that particular design (partially because of the location of the spars, 25% & 70%), not that Cp reached 50% of chord, Cp did not reach back that far.

    The main point being for static testing is that there are some load conditions where the classic loading at stall is not going to check everything. If you're design is like a Lancair 4 or Glassair III you are hitting some airspeeds that result in very low Cl which is actually what moves the Cp aft in this speed range, I'm talking in the range of 300 kts (this would be roughly VNE testing for these airframes) or about Mach .5 . At this mach number for typical 12 or 15% thick airfoils there are no problems from shock waves disrupting the airflow because you haven't gone fast enough yet. Yes, at much higher mach Cp will start to be affected.

    You mention that trim takes care of Cp shifting aft. Trim allows the pilot to ignor shifting Cp by keeping the aircraft balanced and control loads down but even when you trim, the Cp is still at the same location on the main wing airfoil, Cp movement is purely a function of Cl. The structure still has to accommodate the loads that are generated.

    I definitely have to consider your use of a pinned rear spar. I have normally only considered my wing design with the rear spar carrying the bending moment as well as shear. I would think this would make a stiffer wing? My design for a 200+ knot aircraft, wouldn't that be a benefit? What do you consider the pros and cons to be?

    The P-38 tuck under is interesting. Here' how I understand it. The main wing airfoil had it's max thickness further forward than most of today's high speed airfoils (P-38 around 30% chord?) and it was thick enough that when it hit the very high speeds in a dive, the shock wave started at that thickest portion, tripping the airflow into separation. The wing went from the lift it needed to much less lift as the airflow behind this shock wave did not contribute to lift. With the main wing making much less left the nose tucked into a steeper dive. Then the problem got worse. The horizontal tail was a thinner airfoil and didn't form the shock wave as soon but when the aircraft hit a high enough speed that shock wave was ahead of the elevator hinge line causing the control problems you have described. Not fun!

    Thanks for your thoughts!

    Dan Nicoson
     
  11. Mar 22, 2003 #11

    dannicoson

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    Holden,

    I commend you for trying to design a safer aircraft to take care of those that encounter ice inadvertantly. I live in Erie, PA and we have serious ice here most of the winter. This year we didn't get much light plane flying in.

    Concerning the flying slab tail. I know on the heavy iron they work great. The A-6E Intruder had them an it was a great ride at high speed, worked great. My concern in your post is that the anti-servo tab you plan to use is in the same location that the elevator is on the other style tail. Won't it be affected just the same when the airflow goes to h%ll?

    Even a fully slab tail like we had on the A-6E (no anti-servo tab, hydraulically moved) will be useless when you get enough ice on it. Once the ice builds up and forms all those strange shapes, horns and ridges, there's no guarantee any flight will end safely. We rarely worried about ice because we thought we could always power up to a layer with no ice. And that was actually a scary thought process looking back because if we had gotten stuck in busy airspace on a long approach we could have changed shape the same a s a C-172.

    My other big concern with ice is the weight gain. Clear ice often maintains the shape of the airfoil pretty well but it can add a huge amount of weight in a very few minutes. A friend of mine was on a night IFR flight specifically staying out of the clouds. At one point the clouds didn't cooperate. He immediately lost visibility on entering the cloud, ice on the windshield. He called the controller and turned for the nearest airport, 10 miles away, got clearance and made a full power no-flap landing. He said later it was all he could do at full power to get to the runway because of the weight the aircraft gained in that 10 miles (C-172, 100 kts). He said the airfoil shape on the flying surfaces were pretty decent looking but there was over 1/2" of ice on most of the aircraft. Lots of weight.

    I hope you make some progress making simple aircraft safer if they hit ice because it was a long winter without much flying (I can't afford the heavy iron). What kind of testing do you have in mind for qualifying ice tolerance?

    I look forward to hearing of your progress.

    Thanks,

    Dan Nicoson
     
  12. Mar 22, 2003 #12

    orion

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    Hi again;

    Holden, you are right on with your discussion of the next generation of personal jets and layman pilots. We have already seen a number of accidents, which relate to poor piloting technique combined with poor design. But then as the old saying goes, some people have more money than common sense.

    You're also right on with your discussion of the "forward fuselge side wings". These strakes are actually designed for powerful vortex generation and as you indicated, are there for generating huge amounts of lift. Vortex lift is a powerful phenomenon and is used by fighters primarily for aiding them with battle maneuvering. An F-16 for instance, generates over 70% of its lifting force at high aoa on the strakes, not on the wing.

    When I was the configuration designer for the Boeing ATF program, we used similar devices for the program's entry into the competition (this was early on, prior to the later teaming arrangements that resulted in the current monstrocity). The strakes were used for a number of reasons including controling air flow prior to engine inlet ingestion, high lift for combat maneuvering and Cp control for adverse flight conditions icluding icing and when flying after encountering battle damage.

    The trouble with strakes however is that their characteristics are relatively unpredictable, especially toward the regime of high aoa. At times the CL vs. aoa plots look like abstract art - for this reason I tend not to use them in that particular fashion in developing general aviation products. Too many unknowns and of course most private airplane ventures don't have sufficient funding to go through that level of detail in the analysis.

    I've been flying since 1976 and in all that time I've gotten ice only twice. No, I don't fly IFR (at least not by choice) generally and both times of ice accumulation it was in nearly clear air. A short descent was all that was needed to clear it but I can certainly understand the concern. For a short time I worked for a company called Electroimpact, who was developing an ice control method based on electomagnetic repulsion. The concept worked great, even for small accumulations however after I left I lost touch with tem and am not sure if they ever got the sytem certified.

    Dan, to answer your question, you're right, fixing an aft spar would give you a stiffer wing, the question however is whether that is actually what you want. The drawback to this approach can be that you will get stresses in the wing where you may not necessarily want them.

    The Glasair is an excellent example. Due to their continuous structural design (one piece wing), the aft spar is configured for taking some of the bending load. For most flight this is OK, although there is a stress concetration at the aft spar in the area of the wheel well cutout. It turns out that adding material to take care of the stress concetration actualy stiffens that part of the wing and actually makes the stress case even worse.

    At high angles of attack however, this stress concentration becomes the highest stressed part of the wing. If we examine the Part 23 requirements for aerobatic flight and composite structures, by this definition the Glasair wing should not be considered aerobatic. The FARs require 6 Gs limit and a safety factor of 2.0 for composite structures (not 1.5). As such, the ultimate load needs to be 12 Gs. The FEA analysis I did for this condition resulted in stresses well over the failure criteria for fiberglass. This was quite a while ago but if I recall correctly, the ultimate strength of the wing at high aoa was less than 9 Gs.

    For most flight this is OK however, when you add in the long term effects of fatigue loading, Glasairs that are continuously used for aerobatic flight may develop weaknesses down the line.

    For this reason and a couple of others, I prefer to pin the aft spar and let it carry only the local loads of the ailerons and the flaps.

    As far as a stiffer structure is concerned, how much stiffer do you want it? You can design the skin laminate in such a way that it makes up for the lack of a fixed aft spar, and develp the same characteritics without the added complexity of loading up the aft part of your structure. But of course that's up to you - you may have different requirements or reasons than I use in my selections. A stiffer wing is good, especially considering controll crispness and flutter - all I'm saying is that there are different ways to skin a cat.

    Bill
     
  13. Mar 22, 2003 #13

    dannicoson

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    I'll start a new thread

    Bill,

    I'm going to start a new thread "Wing Structure Configuration" because I have a few questions.

    Dan
     
  14. Feb 17, 2004 #14

    StRaNgEdAyS

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    Just thought I'd bump this up again it's got some really good points in it.
    This thread is probably one of the reasons for my interest in the all flying horizontal stabilator. One of the key issues I have to consider is the intended cruise speed for this bird is about 375mph.
    It also goes into some of the thoughts I had which led to the large strakes I have incorporated into my design.
     

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