Skip to comments.Boeing again delays initial 787 test flight
Posted on 06/23/2009 8:46:35 AM PDT by AngelesCrestHighway
Before the markets opened Tuesday morning, Boeing issued a shock announcement that the first flight of the 787 Dreamliner has been postponed again. The company cited a structural defect prompting "a need to reinforce an area within the side-of-body section of the aircraft." Though Boeing chief executive Scott Carson is quoted in the statement saying that "structural modifications like these are not uncommon in the development of new airplanes," the issue appears serious. Adding to the impact of the delay is uncertainty: Boeing said it will be "several weeks" before it will even come up with a new schedule.
(Excerpt) Read more at seattletimes.nwsource.com ...
As a pilot friend always quotes...”better being down here wishing you were up there,....than being up there wishing you were down here!!!”
And the author has *what* evidence to back this up? Sounds like he's planning to short Boeing stock.
I’m starting to have a little nagging concern way back in my mind about graphite airplanes ...
The 787 program is full of usless “managers”. Back 20 to 30 years ago when the 757, 767, etc. were being designed, project and program managers had engineering credentials and years of aviation experience.
Today, Boeing is full of program and project managers that have ZERO engineering or aviation experience but have that all important 6 week course for project manager certification or a cute resume of “consulting”.
It all went from what you can do to who you know and who you blow. Corporate yuppies at their worst.
Serves them right, that is what they get for building 70% of the plane OVERSEAS!
And too much plastic, way too much plastic.(composite)
I already know all the advantages of composites from racing experience. I just do not like its failure scenarios when it does go kabooey.
I wonder if they are installing some “reinforcements” with concerns about the Airbus unkown crash causes in the Atlantic?
65% of plane is manufactured across globe
Sure hope all involved used same units
The areas affected are where the wing made by Mitsubishi in Japan is joined to a part of the center fuselage body made by Fuji, also of Japan.
Yeah, like I would ever be able to afford one of those seats. Lets see a picture of last class.
During a tour of the plant that question was brought up. Boeing’s answer was: “We give your country jobs so your airlines should buy our planes” ... and they do .... the downside of that is those jobs should be here ... BOTOH ....
This announcement is terrible news for Boeing and the US aircraft industry. It is unbelievable that a major flaw was not detected until shortly before the expected first flight. This failure may jeopardize the entire program.
How do they know?
“And too much plastic, way too much plastic.(composite)”
Many airplane designes use it and it works perfectly. The B-2, F-117, Diamond aircraft, and many others all use it without any failures.
Hey, don’t knock the MBAs. They know how to run any business. Just ask them.
(When talking about MBAs, the word “run” is always followed by “into the ground.”)
Mediocre But Arrogant.
Noteworthy Aerospace Ping.
I’m not at all surprised in the delay of first flight, but I am stunned it is because of a structural issue. I wonder if something major broke on the fatigue airframe?
Up until this announcement, Boeing has been running Gauntlet tests on Z-001, fully expecting first flight to be accomplished befure June 30 (which was Boeing’s stated target if first flight “in first quarter 2009.” June 30 is the last day of the first quarter.)
Only a fool would be shocked by this.
This is bad news for employment in Seattle.
It’s the third major delay in the 787 as I recall.
The only good news is that no one is buying airplanes at the moment, so Boeing still has a few months to get it right.
Better check again. Up until Airbus started using it for structrual items almost all other uses in aircraft were for control surfaces and non-structural areas. Composite items offer great advantages but in the real world of using for major areas such as fuselages in every day commercial airliners borders on stupidity.
Get Lego to build it.
I used to work with Boeing on the defense side. One of their hallmarks is coming up within insanely optimistic schedules and then refuse to acknowledge how unrealistic they are. Only at the last minute do they admit they cannot do it. At which point they blame the Government and follow it up with expensive Engineering Change Proposals (ECPs). Other contractors do similar things, but Boeing did it without shame or restraint.
Maybe Boeing needs to run more commercials emphasizing the multi-ethnic makeup of the design team of the Dreamliner. Including handi-capped engineers.
Again? It is going to be like 2 years late just for first flight. They orignally wanted to fly it on 7/8/07.
“Better check again.”
No, you had better check again. I fly a Diamond DA-40, which is a composite.
No kidding. The hallways are not loaded with “make a better product better” slogans. Nope, Boeing has “Diversity is our Strength” slogans. That is the actual slogan they advertise.
Composites don’t suffer from fatigue failures and are stronger than their metal counterparts.
The failures of the 787 are related to metal on composite joints, not the composites.
Some dumb people only hear the word “plastic” and fail to understand that Carbon Fiber is 100% pure carbon.
First of all...I am not an aeronautical engineer.May I assume that the same is true of you? If so,I'll continue.There's nothing in this piece that suggests that the author is an aeronautical engineer.So,I assume,we have two civilians (you and me) analyzing a piece written by a similarly inexperienced individual.
So....how do we know that this is a major flaw? "Major" meaning a serious design/manufacturing flaw that is not easily fixed?
The only individual mentioned in the piece who may well *be* an aeronautical engineer (the Boeing official) seems to think this is easily and quickly fixable.
Again...I think this piece is as likely to involve an effort by the author to violate SEC regulations than it does a serious,difficult-to-fix problem in the aircraft.
I am not an engineer either. My pessimism is based on the schedule delay. It will be several weeks before a new schedule is even announced. Even if the flaw is not substantial, it is unnerving to Boeing customers.
When I worked for GE Aircraft Engines, the slogan seemed to be "Ya' can't see it from 30,000 feet."
Cute little bird ya fly there. The difference between that and an airliners is like a sailboat and the Titanic though.
I am a 3,000 hr pilot up to multi twins as well as an A&E and if you knew the difficulties in inspection procedures of plastic airplanes you would know what I am talking about.
Delamination as well as structural attachment difficulty are prime examples.
It will be one thing for a little single DA40 or any homebuilt composite to have a failure, kill two people and nobody cares. But kill 300, 400+ in a Dreamliner or an Airbus and people care.
Boeing bet the farm on the 747 and won big time. They are also “all in” on the Dreamliner and may have bought the big one.
PS - All the engineers in the world will do you no good if they they don’t know what they are doing and think they are smarter than the ones that came before them.
“and if you knew the difficulties in inspection procedures of plastic airplanes you would know what I am talking about.
My little “sailboat” has a +4.4g/-1.76g limit. Hardly a “sailboat”. Your assumption that just because it crashing will only kill 4 people and not 400 and therefore the engineering is meaningless or less complicated is just plain ignorant and dumb.
Your arrogance is showing. There is a big difference between a mechanic and an engineer. I am an aerospace engineer, you are a mechanic. Big difference in education and understanding of materials. I know how to make them work. You follow my directions on how to maintain them. Don’t pretend to lecture me on materials.
Once again, we see the arrogance of someone wanting to make statments about things they do not understand. This is why this great nation is losing its grip: Arrogance. If you want to play expert then go get the needed education and experience to be one. Until then, just shut up and stop trying to be what you’re not.
There is a slight difference between a DA-40 and a airliner. Especially the pressurized cabin causes a lot of stress to the fuselage structure. The 787 has a cabine pressure level of 6000 ft. That’s much more then the usual 8000 ft. This adds even more stress to the fuselage. This is also a reason why composite military planes usually operate with a much lower cabin pressure of up to 10000 ft. Military planes are also operated less frequent during their entire lifespan.
I hope we won’t see this causing fatal crashes, but i belive we will see numerous incidents caused by fatigue of composite material when the planes get older.
Everey materials suffers from fatigue. You’re absolutely right that carbon fibre is much stronger then the usual aluminium. But we have many decades of experience on long-term aluminium usage, something we don’t have with carbon fibre.
.....something we dont have with carbon fibre......
ASTM has test methods
Sandwich panel flexure testing
he composite materials testing community is moving toward the use of the terms long beam flexure and short beam flexure when addressing sandwich panel testing. The former is used to determine facesheet properties and the latter to determine core shear properties. Such a distinction is logical since we know that, for a...
Column from: High Performance Composites, Contributed by: Dr. Donald Adams
Article Date: 11/1/2006
he composite materials testing community is moving toward the use of the terms long beam flexure and short beam flexure when addressing sandwich panel testing. The former is used to determine facesheet properties and the latter to determine core shear properties. Such a distinction is logical since we know that, for a given applied loading, the flexural stresses (tensile and compressive) in the facesheets increase as beam length increases, but the shear stresses in the core do not. That is, long beams produce high bending stresses while short span lengths do not.
Thus, as noted in my previous column (HPC September 2006, p. 9), a proposed new ASTM standard, “Facing Properties of Sandwich Constructions by Long Beam Flexure,”even contains “long beam”in the title. Conversely, the revision of ASTM C 393, “Core Shear Properties of Sandwich Constructions by Beam Flexure,”while not including the term “short beam”in the title, does refer to the specimen as a short beam in the text.
Despite the differences in nomenclature, it is possible to use the same test fixture for both types of tests. This is because in both cases the support and the loading of the specimen is by means of 25-mm/1-inch wide flat pads that allow free rotation of the specimen while it is being loaded. The commonality between the two test methods is critical. One note before I continue: The term “facing”is used in the various ASTM standards, including in the above ASTM title, whereas HPC uses the term “facesheet,”as in the present (and previous) column, while I have used “face sheet”in my prior works. All of these terms refer to the surface layers of the sandwich panel.
The ASTM documents define “standard”sandwich panel specimen and test configurations, but allow “nonstandard”configurations. The latter recognize the fact that other well-established configurations also exist, and their use may be justified in certain situations. In fact, ASTM recognizes that a nonstandard configuration may even be required if the standard configuration does not meet certain criteria, as will be discussed here.
The standard ASTM configuration for facesheet properties determinations specifies a fixture support span of 560 mm/22 inches and a four-point loading span of 100 mm/4 inches. The corresponding “standard”specimen is 600 mm/24 inches long, allowing some overhang at each support. This is definitely a long beam. The specimen is 75 mm/3 inches wide, thus requiring test fixture support and loading flats capable of accommodating a specimen at least this wide.
n contrast, the standard test for core shear properties determinations requires a three-point loading configuration, with only a 150-mm/6-inch support span. This “short beam”specimen is 200 mm/8 inches long, but is still 75 mm/3 inches wide. Therefore, if a single test fixture is to be used for all sandwich panel testing, it must have a minimum width of 75 mm/3 inches; three-point and four-point loading capability; and either fixed spans of the dimensions noted above or adjustable spans. Fortunately, this is not too difficult to achieve. Figure 1 shows a fixture that meets these conditions. It is shown configured for four-point loading, but when three-point loading is required, one loading head can be removed and the remaining head can be moved to the center of the loading beam. This particular fixture has a maximum support span of 610 mm/24 inches, which is more than required by the standard tests described above, but the full span may be required for nonstandard tests. Since both the loading and support spans are fully adjustable, this fixture can readily accommodate other nonstandard test configurations as well.
Figure 2 is a close up of one of the loading pads on this fixture (the support pads and loading pads are identical). The hardened steel flat pad is free to rotate on the hardened steel cylinder, which permits free rotation of the specimen when the specimen deflects under load, as required in the ASTM standards discussed above.
A noteworthy feature of the ASTM standard test configuration for facesheet properties determination is the four-point loading span of 100 mm/4 inches. For flexure testing of solid laminates rather than sandwich panels (see HPC November 2005, p. 9), the loading span is typically either 1/2 (termed quarter-point loading) or 1/3 (termed third-point loading) the length of the support span. In contrast, the loading span for the sandwich panel long-beam flexure test method discussed here is only 1/5.5 of the support span. That is, the two loading points are considerably closer together than for either quarter-point or third-point loading. In the November article, I discussed in detail the advantages of moving the loading points closer together when conducting any four-point loading flexure test, and the standard ASTM sandwich panel facesheet properties test configuration incorporates this favorable configuration.
In contrast to flexure-testing sandwich panels, when testing solid laminates the support and loading cylinders usually have relatively small diameters, i.e., 6 mm to 10 mm (0.25 inch to 0.39 inch). As discussed above, sandwich specimens are typically supported and loaded by means of 25-mm/1-inch wide flat plates. While the ASTM standards permit the use of 25-mm/1-inch diameter steel cylinders, it is noted that there is a greater risk of local specimen crushing because of the more concentrated loading induced by a cylinder. Any local crushing of the core under a facesheet, particularly the facesheet that is on the compression surface of the beam, is always a concern no matter which loading and support configurations are used. A locally deformed facesheet on the compression surface of a flexure specimen could fail prematurely by local bending or buckling. For this reason, the ASTM standards for sandwich panel testing specify not only flat support and loading surfaces but 3-mm/0.125-inch thick rubber pads between the support and loading flats and the specimen as well. This further relieves local stress concentrations and, thus, reduces the occurrence of local facesheet damage.
Finally, I noted earlier that ASTM defines standard specimen and loading configurations, but also permits nonstandard configurations. In fact, a nonstandard configuration could be required for flexure testing when the particular sandwich panel â short beam or long beam â does not meet certain criteria. These criteria include the required span length, core shear strength and core compression strength. Each criterion is defined as a function of the facesheet’s expected ultimate strength, facesheet thickness, core thickness, core shear allowable strength, core compression allowable strength and even the width of the loading pads to be used. Appropriate formulas for evaluating these criteria are given in the standards.
The ASTM sandwich panel testing standards discussed here will be released when the current round of revisions is complete â a process I expect will take six months to one year. Users of the test methods should find them much easier to interpret and implement than methods defined in previous standards.
Oh reallllllly. I can see you have had an engineeringdectamy(loss of all common sense). The same thing my son has had. Talk about an egotistical ass. Oh, I guess you are one of the engineers that is going to take 20 years and a trillion dollars to do what my generation did in 9 years, go to the moon. I do know what I am talking about sonny, so run along and play with your plastic airplanes.
“but i belive we will see numerous incidents caused by fatigue of composite material when the planes get older.”
We already know the lifespan of composite materials based on the fact there are pressure vessels that use them and have for several decades. So far, no issues. Those pressure vessels hold far greater pressures and cycle more than an airliner.
The pressure at 6,000 feet is a mere average of 8psi difference from the outside pressure. That may fatigue metals like aluminum but composite materials suffer no lifetime degradation.
As we speak there are continual tests being performed on composite materials to ensure they can withstand far greater stresses and cycles than an airliner could ever hope to achieve. They strike them with various metals and forces to ensure accidents cant cause premature failure. They subject them with high voltage and soak them in acid water with many times greater strength than acid rain could ever produce, including that from volcanoes. They purposefully build them improperly to ensure even poorly constructed layers of strands do not cause failure. By far and large, composite materials perform better and are more reliable than their aluminum and steel counterparts.
Toss in the real-world use of these materials from lightweight general aviation aircraft to high performance military aircraft to even space vehicles and warhead shrouds, and we know how they perform. Failure is not expected.
“to do what my generation did in 9 years, go to the moon”
Nothing like admitting you are old and frail and have failed to advance your knowledge. Thank God your generation didn’t follow your lead and stick with stone tablets and loin cloths.
Oh, and for the record, it IS the 60’s generation using these composite materials. They knew about them then and have worked for decades to perfect them for our use now. They did a good job.
“But we have many decades of experience on long-term aluminium usage, something we dont have with carbon fibre.”
As I pointed out in the other post, these are not new materials. They are in common use.
They have been used in submarines, airplanes, pressurized water tanks, acid tanks, tools, space vehicles, etc. Even my 8 pound sledge hammer has a carbon fiber handle. It worked great breaking up concrete. I even use it in high temp applications on my race bikes. Again, this stuff has been around along time now and has earned its place in engineering. I have literally trusted my life to it.
Yep, and the smart ones knew where and what to use it for, non-critical items like race car chassis, ailerons, rudders, cover hatches, etc. Not stuff that will kill 400, maybe 800 people.
Boeing and the rest of the aircraft industry have invested billions and billions of dollars on composites. They will spend whatever it takes and blame anybody else but themselves to save themselves when these composites start taking lives. Oh, I forgot, they already have, Airbus blamed that pilot for using the rudder, tearing off the composite vertical stabilizer and killing all onboard.
And now, Airbus is trying to blame pitot tubes with no evidence.
Composites are nice in a perfect world.
” tearing off the composite vertical stabilizer “
Did the composite fail or did the connecting structure fail? Bad engineering versus bad materials.
Update in Airbus flight manual: Don’t use rudder while in flight.
They recovered the vertical stabilizer intact on this last accident over the Atlantic, rudder and all.
New Dreamliner flight manual update: Do not pressurize cabin during flight operations.
All silliness aside, I will never fly in a pressurized airplane whose fuselage is manufactured with composites. Nor will I allow, if possible, anybody in my family. Wings included. As I said, in a perfect world it is great stuff. But in the real world of commercial aviation, 15 years down the road, sold two or three times and maintained by Mexicans at some Mexican facility who are not even licensed A&E’s glueing a crack in a composite fuselage with 5 minute epoxy. Similar things are happening right now, just not composite fuselages yet.
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