mechanicsupport.com

Unsafe Engineering - Use of NPT Ports in Critical Aircraft Systems

2012-01-28T07:40:00.000-08:00

Axial Crack in NPT Port - Aircraft Fuel Pump

"Extreme care shall be taken when tightening pipe fittings. Overtightening causes distortion, cracking, and leaks." Dept. of Army TM-1-1500-204-23-2 Technical Manual, Aviation Maintenance

NPT Advantages:

Cheap

NPT Disadvantages:

Cracked Ports
Leaking Connections
Not Suitable for Make-and-Break Applications
System contamination due to requirement to use "pipe dope"
Inadequate or incorrect tightening process in maintenance manuals
Difficult to specify a tightening process (operator feel is sometimes required, especially on make-and-break connections that have seen multiple assemblies).
FAA's improper insistence on using a torque value for tightening rather than the design method included in the NPT specification.

I'll throw another one into the pile - poor quality threads. Even in the AN series of fittings the threads quality varies to the point that many will not pass inspection.

The responsibility for NPT connection failure should rest on the shoulders of engineering for specifying such an inferior system rather than the mechanic trying to make it work. "Lipstick on a Pig" When a NPT failure occurs on a critical aircraft system, the first question that should be asked is why was NPT specified?

Pipe Dope Contamination of Aircraft Fuel System

The real problem goes back to engineering, best stated by the United Kingdom Air Accidents Investigative Branch: "The use of any kind of jointing compound at any fuel line connection is fundementally unwise." SAFETY RECOMMENDATION - 2004-010 www.aaib.gov.uk.

Pipe Tape Contamination of Aircraft Instrument System

NPT Spiral Leak Path

Leakage path through threads shown at red points. No matter how tight you make NPT threads, a leakage path still exists. It is the function of the jointing compound to block the path between the crest of the male and female thread.

The use of NPT on thin walled ports is particularly evil as the risk of failure rises:
1. Thin wall ports are weaker and are easier to crack.
2. Thin wall ports strain more (expand from tightening) than thicker walled ports. This expansion reduces the amount of tightening torque required. Weaker port + less tightening force felt = higher risk of failure

At least the engineer can specify NPTF for better safety (quality control) -- but this is seldom done! Better yet is a straight threaded port.

AN Nut Face Design

2012-01-12T07:47:00.000-08:00

AN315 nut showing washer face

The previous article I discussed the purpose of the washer face on the bolt head. AN nuts may also (optional for most nuts) use a washer face. It seems obvious that the seating face should be flat, smooth, and perpendicular (within 2 degrees) to the bolt axis.

So how is this design feature of importance to the mechanic? A lot rides on the integrety of this surface:

Face surface influences the tightening tension produced by your torque wrench by controlling friction. It's estimated that 50% of the friction produced during tightening comes from the bolt head and nut surfaces rather than at the threads.

Nut face angularity influences the fatigue life of the bolt. Angularity or lack of, dramatically reduces the bolt's fatigue life. Bolt Science shows that at 2% angularity (at the edge of allowed limit for a common AN nut) can reduce the bolt's fatigue life from 180,000 fatigue cycles to just 10,000 fatigue cycles!

Nut face surface influences how much embedment relaxation (joint loosening) occurs after you stop torquing the joint. The surfaces squish together slightly resulting in a loss of joint tension. An example of this is the crushed washer below. As the washer surface compresses, the joint becomes loose even though it was properly torqued to begin with.

Now take a look at the nuts below. The malformed washer faces can crush just like the washer and result in a loose joint.

AN310 Castle Nut with malformed washer face

Compare the contact area of a normal washer face on the left with this malformed washer face on the right. The malformed face concentrates the tension onto a small ring which is more likely going to gouge and crush into the contact surface.

Contact face comparison

Deep ridges on the washer face caused by a dull cutting tool

The malformed nuts shown here reduce the joint integrity which may lead to joint failure, which in the aircraft industry is often catastrophic. As I have observed over the years, if the joint fails due to loosening, it will be assumed without any further investigation that the mechanic didn't tighten the joint properly. example, YOU will be blamed. A quick inspection of surfaces: bolt, nut, faying, washers, is recommended before assembly.

and while I am giving recommendations, torque specifications that specify the application of lubricant are defective if they do not also include where to apply the lubricant. Since the seating surfaces represents 50% of the friction during tightening, it's important to know if the engineer intended that these surfaces, and the thread surfaces, be lubricated. The practice varies among engineer's so there is know way of knowing if the specification doesn't state.

Companion article link AN Bolt Head Design

AN Bolt Head Design

2012-01-08T06:24:00.000-08:00

Dear Sir/Madam,

Some 12 months ago I was asked, 'what is the reason for the machined section under the head of some bolts' ?. This I should have know as a licensed aircraft maintenance engineer for more than 40 years, yet the question took me by surprise because I never had cause to question the problem.

Today I'm in my 94th year and all of my working life has been in the Automobile and Aircraft fields, of course the Department of Civil Aviation (now CASA) has tested all engineers involved that they know much about bolts, and in particular AN bolts (Air Force/Navy). We must know of metalurgy, tensile strengths (UTS), yield points, cadmium plating, and the dangers of chrome-plating bolts the requred knowledge seems endless to the stage where BOLTS as such is a complex science, and this leads me to writing this message.

In all my years working with bolts I would question most of them, especially in aircraft-but in all those years I never asked why many bolts today have a slightly raised circular section under the head of the bolt within that part of the imagined circle within the under side of the hexagon flats. The raised machine section is only a few thousandths of an inch proud, in the order varying about : 004" to :008" in the ones I've measured.

To answer the question I said I didn't know! adding that I had never been asked, nor had queried the reason for the raised section, further adding that I'd find out!

With the lack of better knowledge, I suggested to the question that the raised/proud section might be to ensure the first 'bite' in tensioning a bolt fastener is centered immediately close to the bolt shank and that increased tensioning would gradually spread outward from the bolt shank thus ensuring that the initial axial loading would essentially take place radially over the raised portion of the under-head hex, with or without a washer.

To me it makes a lot of sense to have this section machined thus preventing sharp edges from the hex edges gouging into the clamped pieces, yet of this
I'm unsure.

AN bolt showing washer face

This area is called a "washer face" and defines the bearing area for the bolt head. The bearing area is useful for calculating bearing loads on the washer and/or faying surface so you do not exceed the material's yield strength (crush the joint or washer). It also provides a flat machined bearing surface.

Crushed washer from tightening beyond material's yield strength

Washers resting on washer face

I would guess that it is easier to achieve a specified bearing area by machining a circle than by beveling the edges of the hex nut. Another alternative is this:

Here is a link to an article I wrote a while back that touches on the subject of washer unimpressive strength.

Companion article link: AN Nut Head Design

B-nut Torque and Loosening

2012-01-01T07:53:00.000-08:00

"Failure of maintenance personnel to properly tighten the fuel supply hose at the engine-driven fuel pump."
Injuries: 3 Fatal.
NTSB Identification: ERA09FA068

It's long been recognized in engineering and among the common man that properly tightened threaded fasteners can become loose. There is an entire industry devoted to making devices that prevent properly tightened threaded fasteners from working loose; lock nuts, lock washers, adhesives of many types, special thread forms.Yet, none of these devices are used on aircraft B-nut connections. (A few aircraft did use lockwire drilled B-nuts, but these are seldom seen).

A survey of aircraft accidents where the B-nut was found loose reveals a bias among accident investigators. Investigators are not investigating why B-nuts may be loose because their bias tells them they already know why; the mechanic didn't tighten it properly. This sloppy and unprofessional work degrades the entire process of accident investigation. Any non-retained threaded fastener can work loose. It is not acceptable to assume that any loose threaded fastener was caused by " improper torque" by the installer without doing some additional inspection work. There are many reasons why a B-nut can work itself loose: thermal expansion and contraction of the joint, vibration, malformed seating surfaces, etc.

A pilot or mechanic or engineer who fails to learn and repeats a mistake can be the probable cause of a future accident; an accident investigator can also be the probable cause of an accident as he had it in his power to prevent it but did not.

The Aircraft Structural Mechanic (why you deserve a pay raise)

2011-12-22T16:31:00.000-08:00

"you spend too much time lookin" was the latest comment from the boss; a reflection of a time long gone when aircraft mechanics were first fixit men, and then parts changers. Aircraft Mechanic's lookin is what keeps the airlines flying; it's how transport class airplanes are designed; it's designed into the structure; easy access for lookin, lookin tools, lookin techniques, and methods. Does the organization understand this? If not -- and the aircraft is designed to be damage tolerant --it's not tolerant but a time bomb.

It is expected that the structure overtime will develop damage but that it will be found by the lookin mechanic before it becomes fatal. Damage tolerant structures must have a "high probability of detection" The organization needs to provide the personal, tools, environment to make this possible.

Large aircraft are designed to be damage tolerant 1. -- there is no limit to service life. Aircraft are kept in service by a partnership between the structural designer and the structural mechanic. This is no "remove-and-replace" maintenance; "structural maintenance is the cornerstone for ensuring continued airworthiness of damage tolerant structures." 2.

Damage tolerant allows for cracks in Structurally Significant Items (SSI); it requires timely inspections to detect such damage with a high probability BEFORE residual structural strength falls below specified values. There are no "standard" repairs or inspections. Each inspection and repair to a SSI is designed by the structural engineer, communicated to the structural mechanic without ambiguity, and performed as the engineer designed. When this does not happen, all hell breaks lose and you get Japan 123 type accidents.

"Inspectability" Where and When to Inspect is a key element
Damage tolerant only works when you know where and when to inspect. Fatigue cracking is cumulative with respect to aircraft usage so it is a straight-forward process to monitor. What is not so easy to predict is corrosion damage; both from standard corrosion and stress corrosion. Stress corrosion reduces fatigue life. A damage tolerant structure must include a Comprehensive Corrosion Prevention Program(s). If this program is deficient then fatigue life estimates are not accurate and the whole concept of damage tolerance goes out the window and leaves the structure weaker than anticipated. (Aloha Airlines 243 for example) Corrosion prevention, detection, and removal is required for a damage tolerant structure and the execution of this program is part of the Structural Mechanic's job.

This is the concept and it has worked well. To give an example of just how well: Boeing's 737 had a minimum service design objective of 75,000 flights but high-time aircraft have achieved 90,000 flights. Exceeding the design objective occurs across the Boeing fleet including the 707, 720, 727, 737, 747. 2., 3.

This is not only a Boeing accomplishment but also reflects the performance of the aircraft structural mechanic in meeting the engineering expectations in regards to inspecting, detecting, and repairing. If the organization don't have good lookers then they shouldn't be flying damage tolerant aircraft. You can't attract and keep the best and most responsible aircraft mechanics unless you provide the best wages, benefits, and working conditions.

1.In 1978, the FAA adopted “damage tolerance” as the preferred choice for managing fatigue in civil airliners.
2. Fatigue Issues in Aircraft Maintenance and Repairs, Ulf G. Goranson, Boeing Commercial Airplane Group.
3. Damage Tolerant works only for defects that are detectable. Adhesive bonded structures may have structurally significant defects that are not detectable.

Aircraft Hose Bonding and Lightening Protection

2011-12-21T07:42:00.000-08:00

Lessons from Agusta Bell 206B JetRanger II, G-AWMK

AN919 anodized aluminum fitting is an insulator

Current commercial and military-aircraft standard for electrical bonding: 0.0025 ohm for lightening protection and RF potentials.

Unlike other electrical systems, aircraft systems use the structure (skin and/or airframe) as a current-carrying-conductor. There is no "neutral" wire in aircraft. The aircraft skin and components (and hoses) carry the return current back to the battery.

Aircraft can develop high static electrical charges as is evidenced by the need for static-dischargers. Arching can occur between aircraft parts that are at different electrical potentials. In some aircraft, hose is routed through the fuel-tanks. Arching within the fuel tanks can occur if a bonded hose is within spark distance of an unbonded hose (Augusta Bell 206B had an in-flight fuel-tank-explosion).

Aircraft can be hit by lightening. Bonded components help the lightening current to flow through the airframe without arching. A bonded component is where a electrical conductive path exists between two aircraft parts. A common example is the installation of a bonding-strap between the engine mount and the airframe. A metal braid fuel hose or metal tube that is not bonded may have the potential to create arching or sparking during a lightening strike.

Most aluminum fittings have an insulation layer on their outer surface that prevents electrical bonding. This insulation layer is called "anodizing". Anodizing colors the aluminum (as the picture to the left shows) and protects it from corrosion. But, it is also an insulator. Removing and installing the hose or tube several times will wear through the insulation on the threads and sealing surfaces.

Aircraft Hose Assembly with anodized aluminum fittings. Anodized aluminum is an insulator.

In the case of the Bell 206B fue- tank-explosion, the hose was not bonded because of the anodized coating and this led to a static discharge between the unbonded hose and a nearby bonded hose. If your application requires bonding, then be careful with anodized aluminum aircraft parts. They are insulators.

Aircraft Teflon hoses (those meeting mil-specifications) have a conductive layer of carbon black to provide provide electrical conductivity and prevent static charges. Commercial (non-aerospace) hose does not have this static control.

Aircraft Teflon hose assembly showing carbon black inner-liner

Military Specification MIL-H-25579E requires that hoses (through -8) be capable of conducting a direct current equal to or greater than 6 microamperes with a test potential of 1,000 volts dc between the hose inner liner and one end fitting. This prevents the build-up of static charge and arc pin hole leaks to the wire braid.

When teflon hose was first used on aircraft it developed pin-hole leaks. The plastic Teflon develops a static charge so great that it arcs to the grounded steel braid causing a small hole in the Teflon.

--story time--

The classic example of bonding is the pilot who complained that every time he started his airplane the mixture control knob got warm. It turns out that his engine mount wasn't grounded to the airframe. The only conductive path for the battery current was back through the mixture control cable. I have also seen this happen with a metal braid oil pressure hose going from the engine to the gauge. All the starter current flowed through the hose braid. The braid lit up like a heater element and cooked the hose. Fortunately it was oil-pressure and not fuel-pressure.

Take a look at that fat #2 or #4 battery cable going to the starter. All of the current flowing from the battery to the starter must also flow back to the battery. It does this through the airframe. The current will take the path of least resistance. Good electrical bonds help deliver current to the starter for quick starts. They also help to keep the current out of your instruments and hoses where it doesn't belong.

Suspect Un-Airworthiness

2011-11-28T07:06:00.000-08:00

Letter received from a friend:

The pic is of a freshly topped XYZ engine. We sent it back to XYZ cause we had metal in the filter and because they had fitted the wrong rockers. This engine is on it's third camshaft since last overhaul. Pilot is not happy with the vibration and missing. Pull filter and there are 2 strings of red silastic and metal. As you can see in the pic, they have put silastic between the case and the cylinder base. Suprisingly they have had a rash of through bolt and stud failures. Can't imagine why. They just don't see a problem. You'll love this; they rang and said that at oil change time should be pretty normal to find a match head of metal in the filter, comes from the cam and wear on the cylinder walls. I'm now the baddy in this. If I let it fly and the engine packs up over rough country, I'm in the gun. If I dig my toes in, will they strip it and admit it has something wrong.

So what do you do when you discover serious problems with an engine and the manufacturer's service rep tells you "it's normal" don't worry? "Put it in writing" is my immediate response but beyond that how do you approach the problem? Here is my suggestion.

You can go round-and-round on what is "normal" and what is "airworthy" with manufacturer's and repair agencies but this misses the point. They all work to an approved process that is kept in check by their quality control system. Deviations to the approved process that escape the quality system are evidence of a process deviation and a quality system break-down. Product released during the time of deviation is an escape and is not approved nor airworthy because it does not, or is suspect, as being not in conformance. Part of the definition of "Airworthy" is that the the part conforms to its type design." Airworthiness is not the determination by the companie's service rep that unapproved material in the oil system is normal.

The exact nature of the non-conformance i.e. silastic or bits of metal, is not the issue. The issue is that the engine escaped their quality system, has suspect deviation from their quality process and is therefore "not airworthy" The product can only be re-inspected and made airworthy after the process and QC system is brought back into compliance as evidenced by corrective action and audit to confirm that the "corrective actions" are successful.

It appears that there is a break-down in their quality control system. I would not accept repairs as "approved" or "airworthy" during the time of non-conformance. Can they provide the date of the last quality audit and what corrective actions have been made? I would be concerned that your response was an individual opinion and does not reflect the requirements of their quality system.

Put your concerns in writing. Send it to XYZ with a copy to the owner and whatever you call your "FAA". Any decision is outside of your area of expertise -- you can only express a "concern" Let engineering "XYZ" and standards "FAA" make the decision.

I have seen a Lycoming failure caused by a "match head' size of Silastic that entered the oil system and stuck at the rod bearing oil port where it starved the rod bearing of oil. My recommendation based on experience is to fully inspect all oil passageways for unapproved material.

Normal wear does not occur as metal chunks.

Parting sealant "Silastic" loose in the engine indicates that gasket material has extruded from between the parting surfaces causing cylinder-to-crankcase embedment relaxation. I doubt joint loosening caused by gasket extrusion meets design intent.

----------Another potential action you might take if circumstances apply------------

Since you must determine if the item is airworthy (Airworthy. To determine that the installation of a part complies with the applicable regulations, the installer of the part is ultimately responsible for establishing that the part conforms to its type design and is in a condition for safe operation.). AND you suspect that it might not be then (at least in the US) you can file a SUP if the company is a PAH Production Approval Holder. A Suspected Unapproved Part (SUP) is "new parts that have passed through a PAH’s quality system which do not conform to the approved design/data." AC No: 21-29C Change 1, Definitions p. Unapproved Part(2)

Aircraft Corrosion -- Cabin Condensation Calculator Review

2011-11-25T09:29:00.000-08:00

Review from Mechanic's Toolbox software

How to Properly re-magnetize a magneto rotor magnet?

2011-11-25T06:43:00.000-08:00

Hi,

Just found your sight. Seems like a great resource.

I work on aircraft junk we run on airboats. I like to know all the proper procedures and information I can learn. In what book or lesson can I find out how to properly re magnetize a rotor?

Your help is appreciated.

Gain a thorough understanding of the Magnetic Hysteresis Curve:

Hysteresis Curve for a Ferromagnetic Material

Allowing the magnet to drop from E to F on the curve is the most common mistake.

Here is a link for a pretty good description of the curve

A link to a more detail description I wrote some years back

Aircraft Structural Screws - Video

2011-11-23T15:50:00.000-08:00

Metal Fatigue, Cracks, and Turbo Mallards

2011-11-15T06:40:00.000-08:00

Fatigue Failure with attempted repair

Metal Fatigue occurs when the metal is subjected to repeated or alternating stresses below the material's static yield strength. Fatigue failure occurs BELOW the material's ultimate tensile strength. Parts that are exposed to alternating stress cycles, such as engine crankshafts, may break even though they were never stressed near their ultimate strength. How does one know if a part is close to fatigue failure?

Fractured aircraft engine crankshaft. Beech Marks are a sign that a crack progressed across the part and failure was due to metal fatigue (red arrow). The white arrow shows the crack initiation point.

Fatigue life is determined by the number and magnitude of the stress cycles. Fatigue cannot be inspected -- unless you know the past history of a part, there is no method of determining how many stress cycles and therefore how close to fatigue failure the part is at.

Aircraft components that are described as being "zero timed", "like new", or "restored" are marketing terms that do not describe the remaining fatigue strength. That is the challenge of aging aircraft and purchasing critical stressed components where the past history is not known.

Engineering Critical Assessment:

Without knowing the past loading history, the only method of evaluating the failure potential by fatigue is through an engineering critical assessment using Damage Tolerance methods such as fracture toughness, allowable flaw size; and through this process inspecting for existing flaws and calculating the tolerable flaw size for the projected future loading spectrum. What this means to the mechanic is that maintenance cannot prevent fatigue failure, cannot inspect for fatigue failure, nor determine airworthiness from a metal fatigue basis without something to inspect; and without an Engineering Critical Assessment there is nothing to inspect.

This is the issue that the mechanics of N2969, the Turbo Mallard who's wing broke off killing all 20 people aboard. An old airplane with skin cracks -- where is the point of failure; it could be the moment a crack is formed, or it could be a defined crack length based on an appropriate fracture mechanics analysis and following applicable codes. It is safe to say that all aircraft have cracks and that not all cracks are a point of failure; in each case what is appropriate--replacement or repair? Without unambiguous maintenance standards based on engineering analysis, fatigue failures will continue regardless of the intensity or quality of maintenance.

"Old designs are never proven for fatigue simply by virture of their longevity. Fatigue is wear-out. There is no guarantee that future failures will be confined to those seen in the past." Steve Swift, GNATS AND CAMELS - 30 Years of Regulating Structural Fatigue in Light Aircraft"

Comet 1 SN Diagram Animation

Inspecting Aircraft Control Cable Video

2011-11-12T09:51:00.000-08:00

Short video on control cable inspection. Link for full size

Aircraft Rivet Hole Fatigue Strength

2011-10-06T08:24:00.000-07:00

Smoking rivet and failed stop-drilled holes

One lousy hole out of thousands and the aircraft crashes killing all occupants: "The points where the fatigue fracture originated were in a rough drilled surface where the edge of the drill had left a sharp corner at the change in section thickness near the bottom of the hole." Loss of Helo H-295 August 21, 1971

Every open hole distorts (strains) under loading. Cyclic loading results in repeated loading and unloading of the hole. This is the mechanism for fatigue crack initiation and growth. A tightly installed plug (fastener) in a hole inhibits this strain-deformation.

Burrs increase stress concentration at hole edges. Crack-growth is largely independent of material tensile strength.


	Chart from Repairs to Damage Tolerant Aircraft by T. Swift, Federal Aviation Administration, FAA-AIR-90-01.

Proper riveting (filling the hole) increases fatigue life over an unfilled hole. An open hole increases stress by a factor of three times. A filled hole reduces this stress concentration to two times. An interference fit filled hole, for example bucking a rivet into the hole further reduces the stress concentration factor. Per FAA-AIR-90-01 Fig. 20.

Stop-drilled holes should be filled. Also, use proper rivet technique so that the rivet swells and fills the hole. This places the hole boundary in compression. For the rivet to swell and fill the hole the rivet must be driven squarely and not "clinched" When clinching occurs the hole is not properly filled and swelling does not occur. Thus the beneficial residual compressive stresses are not present. When this occurs the fatigue life is no more than an open hole.

Static Strength Considerations

Rivet tear-out

Up till now the discussion concerns cyclic strength. What about static strength? General considerations for the mechanic regarding holes and static strength are as follows:

For isotropic materials (materials like an infinite sheet of aluminum that are the same in all directions):
The presence of a hole has little effect on the static fracture strength of a ductile material.
The presence of a hole has a large effect on the static fracture strength of a brittle material.

For non-isotropic materials such as laminated composits:
Fracture strength is a function of hole size "hole size effect" and for multiple holes, hole interaction - in short the subject is complex.

High MP low RPM Continental TSIO520 engine (or Lycoming engines)

2011-09-19T11:26:00.000-07:00

Hi John. I enjoy and learn a lot reading your monthly mechanic`s tool. Thank you very much.

I am a pilot and would like to your point of view from a discussion some of us been having for quite a time: it is there any problem on a turbocharge engine to operate on low rpm and high manifold press. i.e.: cessna 421 at 40" and 1900 rpm for a 5 to 10 minutes period of time?.

I would appreciate your comments.

Your joking right? Your not aware of Cape Air/Hyannis Air Services Inc.'s engines slinging pendulum absorbers through the crankcase?

http://www.avweb.com/avwebflash/news/Cape_Air_Grounds_Cessna_402_Fleet_195402-1.html

or, Continental Service Bulletin SS107-5,

Or this from Continental:

SUBJECT: MINIMUM CRUISE RPM LIMITS
PURPOSE: To inform operators of the possible long term effects of low engine RPM in cruise conditions. To establish limitation of minimum engine RPM in cruise.
COMPLIANCE Upon issuance of this bulletin
MODELSAFFECTED: O-470-G; IO-470-N; IO-520-BB, CB, MB, P; IO-550-A, B, C, D, E, F, G, L, M, N, P, R; IOF-550-B, C, D, E, F, L, N, P, R;
TSIO-520-AE, BB, BE, CE, DB, EB, JB, KB, LB, NB, UB, VB, WB; LTSIO-520AE; TSIO-550-A, B, C, E, K; TSIOF-550- J;
TSIOL-550-A, B, C
Teledyne Continental Motors (TCM) has examined recent occurrences of crankshaft counterweight release and subsequent engine stoppage in two high time IO-520 and two high time TSIO-520 engine models. Investigation and reported service history lead us to believe that these occurrences are associated with engine operation at sustained cruise engine RPM of less than 2300 RPM. Power settings of less than 2300 RPM have been within the recommended cruise range allowed by TCM’s Model Specifications. It is TCM’s belief that the population of aircraft equipped with the affected engine models that operate using an RPM less than 2300 RPM for extended cruise operation is limited. TCM will continue to evaluate these reported counterweight releases in an attempt to establish a root cause, including any possible connection with power settings. TCM has not been made aware of any additional confirmed occurrences of this type beyond those mentioned above. Effective immediately, TCM strongly recommends the following limitation be observed on all the models affected above:
Engine cruise RPM settings should be no lower than 2300 RPM.
NOTE … This limitation applies only to cruise operation and is not meant to
supersede the aircraft manufacturers’ recommendations for other operational modes such as emergency or holding procedures. Any engine listed in the models affected that has been consistently operated outside the recommended limitation in this bulletin should contact TCM Technical Customer Service at 1-888-826-5465 Option 1 or 1-251-438-3411 x8299 for further information and instructions

Pendulum Absorbers mounted on crankshaft

-----------------------Here is the problem -----------------------

Counterweights are in fact pendulum absorbers that have a fixed capacity to absorb torsional crankshaft energy. If you feed in more energy than they can absorb they "detune" or "jump" . By detuning, the absorber, which is free to swing like a pendulum, no longer swings but bounces around violently. This violent bouncing will break or knock out the retaining rings and plates and detach the absorber. How do you feed in more energy? - by increasing the torsional twisting of the crankshaft. You do this when you:

a. increase engine torque by increasing MP

b. operating at an rpm that coincides with the crankshaft's resonant frequency. The resonant frequency is around 2000 rpm for the 520 crankshaft.

So by operating at low rpm/high mp you are close to peak energy input into the 6th order absorbers. But there is more to this story. When the absorber is far from peak energy it kind of sits there and wears a depression into the bushing (frets). This changes the pendulum length and the absorber's natural frequency. This means that its energy absorbtion capability is reduced. So as you're engine gets to "high-time" or close to tbo in hours, the 6th order counterweights are more sensitive to detuning.

Now about operating for a short period of time? Keep in mind that once the absorber detunes it jumps to a different curve and doesn't come back into tune unless you bring the power back to close to idle. Think of a child on a swing - your absorber is suspended in exactly the same manner (bifler suspension - at two points). The child swings smoothly, but if you disturb the swing it stops swinging and you must grab the child and completely stop the swing and start over.

Crankshaft with pendulum absorber removed.

Circlip, Plates and internal pin provide a bifler attachment

Exhaust Valve Deposits - Concentricity

2011-09-06T07:41:00.000-07:00

Look for deposit concentricity. This pattern can only occur if the temperature across the valve face is the same at any distance along the radius. Even temperature can not occur if the valve is leaking hot exhaust at a spot on the circumference.

Concentric

Concentric - deposits share same center axis shows that valve face temperature is the same at any distance from the center. If the valve face were leaking then the temperature at that spot would be hotter and the deposits would no longer be concentric.

Continental Exhaust Valve Concentric Deposits

Not Concentric

Not Concentric - valve face leakage

Not Concentric

Exhaust valve leakage

Not Concentric

Lycoming exhaust valve leakage

Aircraft Washer Usage

2011-09-02T07:51:00.000-07:00

Rule of Flat Washers:

All washers shall be made from a material which is capable of accepting the peak fastener load without deformation.


Incorrect washer used on NAS148 high tensile strength bolt led to the loss of N76195 and its occupant

A washer can provide multiple functions, the two most important ones are:

1. Spreads the clamping force over a larger area to avoid compressive yielding, and

2. Hard, smooth, consistent material for good preload (clamping) control.

Other functions are to:

1. Prevent galling of the nut face or surface during tightening.

2. Reducing the external load carried by the bolt by increasing the effective pressure area. This stiffens the joint members and the stiffer the joint members the smaller the fraction of external load the bolt will "see".

3. Prevent galvanic corrosion by separating dissimilar metals. Example would be using an aluminum washer under a steel bolt head tightened against an aluminum crankcase. Any galvanic corrosion occurs between the washer and bolt head rather than between the crankcase and bolt head. A washer is cheaper to replace then the crankcase.

4. Increase energy stored in bolt by using a longer bolt. This helps retain clamping force.

5. Adjusting grip length.

Washer Strength:

WASHER compressive strength MUST be matched to the BOLT/NUT combination!

Pictured above is a low-yield strength hardware store washer placed under a propeller bolt. Low-yield strength washers that score/crush in-service under high strength BOLT heads or NUTS relieves the clamping force, eventually resulting in propeller detachment during operation.

High-Strength Aircraft Washers MS20002

MS20002 smaller inside diameter for closer fit to bolt shank

Standard AN960 has larger inside diameter - less bearing surface area

A size comparision of the common AN960 washer with the harder MS20002 reveals that the MS20002 washer has a smaller diameter inner hole and a slightly larger outside diameter. The non-chamfered version offers approximately 20% more surface area to the nut. For example, a 1/4 inch AN960 has a surface area of .13989 sq. in. compared to .16998 sq. in. for the MS20002, thereby reducing the stress per square inch on the washer by spreading the load over a larger surface area. This helps prevent washer or faying surface crushing and reduces joint embedment relaxation. A quick calculation shows that when a AN4 bolt is fully torqued the stress per square inch on the washer reduces from 18,000 psi with the AN960 washer to 15,000 psi for the MS20002 washer

Comparing Thickness - high-strength washer next to head is thicker than standard washer

MS20002C - bevel to clear radius at bolt shank to head

Aircraft Washer Usage Chart

When not to use a washer

Incorrect - no washer needed or desired here

When NOT to use a washer. The built-in washer under the head of a flange head bolt acts to distribute the clamping load over a greater area. No washer is needed or desired. This aircraft starter is assembled with washers under the flanged bolt head. Notice that the bolt head overhangs the washer more on the left side. Bolt now has prying tension.

Calculating Surface Pressure:

Not only the washer but the joint (faying) surfaces must have adequate compressive strength.

Crushing - tightening beyond compressive yield strength of bearing material.

To calculate bearing-stress (surface-pressure), you take the bolt-tension and divide by the contact area between the bolt head and the part. You then compare this value to the allowable surface pressure for the joint material. The allowable contact stress for material is usually about equal to the ultimate tensile strength due to the nature of localized forces on solid bodies.

A rule of thumb is that the allowable surface pressure is approximately equal to the material's ultimate tensile-strength (due to elastic and plastic constraint from the surrounding material). Even if you reach the pressure limit, that just means you begin indentation of the part, which does not necessarily mean part failure. You will need to decide what the part limits are with respect to static and cyclic loading, temperature exposure, etc.

Non-Metallic Washers

Aircraft Spinner with non-metallic washer

Some aircraft applications use a non-metallic washer such as under propeller spinner screws. Typically, these washers are made from high compressive strength phenolic or sometimes nylon so that the screw can be tightened without crushing the washer. Some have suggested using a "Teflon" washer.

PTFE "Teflon" is has a tendecy to creep under compression (cold flow). In other words "it runs away from the stress" and leaves the screw loose. This low compressive strength can result in loose fasteners and joints when used as a washer under screw heads. To illustate the low compressive strength of PTFE, the chart below compares Nylon's compressive strength to PTFE. Care needs to be exercised when substituting materials that the substitute has suitable mechanical properties to function as well as the original.

Plastic washer material strength

Propeller blade cracks

2011-09-01T17:05:00.000-07:00

Some interesting stuff on inspecting propeller blades for cracks before flight. Australian Airworthiness Bulletin 61-008 "To provide guidance on propeller continuing airworthiness/maintenance practices"

A few quotes from the bulletin:

detection of the crack may only be possible from the rear of the blade

I always looked at the front - good advise - article has an good explanation as to why. Another quote:

Investigation has shown that cracks have propagated over a long period, which in some cases exceeds thirty ground/air/ground cycles i.e. thirty flights. There is no evidence to suggest that failures have occurred where a crack may have propagated from initiation to final failure in one ground/air/ground cycle i.e. one flight. Therefore detection of the crack and prevention of failures of this nature should be achievable.

A quick inspection before flight can prevent blade failures.

Some blade paint schemes are not conducive to easy inspection of

the rear surface of the blade...

This is the big point! A thick durable layer of tough epoxy paint might hide the crack!

Piston engines and engine mounts were painted with a thin coat of brittle enamel paint. Through long experience we found that cracks would appear through the paint. The paint did not prevent inspection. Now the customer wants a thick powder-coat gloss finish. Cracks and corrosion are hidden; inspection is hindered or made impossible. Aerospace is more concerned with performance, endurance, inspectability than cosmetics and bright colors under the hood.

Slick Magneto Inspection Tip

2011-07-21T07:23:00.000-07:00

Slick Magneto Arching damage

Engine roughness has many causes. Here is one item to check; quick and easy:

Remove the harness caps and inspect the distributor block. Look for:

Erosion or burning on towers
Color differences in lead contact buttons
Carbon dust

Notice the burning (arching damage) and color change. Lots of erosion; this engine had a miss for a long time.

Slick Magneto Arching damage Closeup

This damage is caused by the electrical arc bypassing the spark plug and finding an easier ground path along the lead tower and to the magneto housing. The magneto below has a different problem.

Slick Magneto Carbon Dust

There is a layer of black carbon dust shown by the red arrows. Carbon dust is conductive and can cause arching to ground inside the magneto instead of at the spark plug. Lets take a look inside this magneto.

Slick magneto carbon brush

This is the worn carbon brush inside the magneto. Just as we suspected from all of the carbon dust on the distributor cap. Here is what internal arching does to the insides of the magneto. That carbon dust is bad stuff.

Slick magneto arching damage

Notice the white residue. You will find this inside the magneto cap. If you remove the harness cover and see white residue on the lead towers then there is lots more inside. In case you're interested; this magneto did run and pass a mag check -- it just crapped out at full power and resulted in an aborted takeoff.

Slick Magneto white residue

There have been improvements in the Slick magneto. The picture below shows the carbon brush and the insulating portion of the distributor block. Notice the "dams" (yellow arrows). They are new as of about 2010.

Slick magneto distributor block closeup showing dams

These "dams" serve the same purpose as the ones below on a electrical transmission tower insulator. You might have also seen these on some spark plugs.

Electrical Tower Insulator

Original (old) style

This older style has no dams. The dams capture the dust and provide a longer electrical path to ground.

Removing the harness cap and inspecting the top of the distributor block can be quite revealing and save you troubleshooting time; it can find problems early and at little cost. Of course all of this stuff I write about here is part of Mechanic's Toolbox Software.

Aircraft Engine Age Deterioration

2011-07-05T15:39:00.000-07:00

In Service Condition Inspection of N6937Y, PA23-250

Compression: Excellent
Oil Consumption: Good
Oil Analysis: Clean
Oil Filter Examination: OK
Engine Operation: Smooth

Based on above findings aircraft was operated on June 23, 1996. Aircraft crashed during flight killing all aboard.

Fatigue crack at corrosion pit between cylinder fins
In flight fire
Wing separated in flight
5 fatalities

Lycoming connecting rod with rust

Overhaul Condition Inspections that WERE NOT DONE but recommended (Factory overhaul inspections after 12 years; engine time-in-service: 21 years):

Corrosion and Pitting Inspection
Magnetic Particle Inspection
Florescent Penetrant Inspection
Eddy Current Inspection
Visual Inspection
Dimensional and run-out inspection
Wear and surface damage

These more thorough inspections were not performed presumably because In Service Condition Inspections were all that were required (in the opinion of some airplane owners and mechanics) to determine the safety of the engine. Unfortunately, In Service Condition Inspections were not adequate given the age of the engine. "Bad practices that result in no immediate ill effects wind up becoming the norm."

Age related deterioration may result in sudden and catastrophic engine failure as this example illustrates. This accident, and other age-related-failures, may be prevented on high calender time engines by using more thorough Overhaul Condition Inspections that better inspect for corrosion pitting, fatigue cracks, and other deterioration both external and internal.

Links to accidents related to corrosion pitting:

Lycoming fuel injector lines

Crankshaft Failure of antique engine

Navion Crankshaft Failure - pitting on fillet

A 150 µm deep (0.010 inch) corrosion pit
A post-mortem examination found evidence of soot in the airway of the pilot, which indicated that he had been breathing during exposure to smoke. Toxicology results showed the presence of cyanide in the pilot’s blood at a significantly elevated level; cyanide is a common combustion product of some materials found in aircraft construction.

Cessna 152 Cylinder Failure

Engine Failure/Fire Piston Pin Pitting

Propeller Loss "Had the failed engine been overhauled within the manufacturer’s recommended time of 2000 hours, or even within 2200 hours had it met the manufacturer’s 200-hour extension requirements, the overhaul would have occurred before the flange cracking had reached a critical stage and the crankshaft should have been scrapped."

The occupants were fatally injured. "A fatigue crack developed in the engine crankshaft as a result of corrosion pitting and the absence of a case-hardened layer on the fillet radius of the number six connecting rod journal. The fatigue failure of this section of the engine crankshaft resulted in a complete loss of power."

Both occupants were fatally injured. The helicopter was completely destroyed in the post-impact fire. Enstrom F-28C Helicopter C-GVQQ total time in service of 611 hours over the 27 years since the 1982 overhaul. "The fracture of the check ball retainer in the exhaust valve hydraulic tappet..."

Calender time and Hours (stress cycles) are both considerations for continued safe operation

I0-520 Cylinder with combustion chamber cracks

"Old designs are never proven for fatigue simply by virture of their longevity. Fatigue is wear-out. There is no guarantee that future failures will be confined to those seen in the past." Steve Swift,GNATS AND CAMELS - 30 Years of Regulating Structural Fatigue in Light Aircraft"

Oil Leaks -- It's not the gasket it's the surface

2011-06-22T06:00:00.000-07:00

The only reason we use gaskets is because we can't machine a truly flat surface. Bugatti engine blocks were hand scraped to ensure that the surfaces were so flat that gaskets were not required for sealing! Almost true - they were hand scraped but Bugatti engines did leak oil but the point is that surface flatness is #1 when it comes to sealing surfaces.

Check your surfaces for flatness. A customer brought this cover plate into my store to purchase a gasket. I have this habit of placing plates on the counter upside down and pressing on the edges to see if they rock - a quick check for flatness - this one rocked! A new gasket wouldn't work any better on this plate than the old one. Pretty amazing to me that the customer was unaware of this. He is going to have to flatten this plate.

Another check for flatness is to lightly lap the surfaces and inspect the lap contact area. When I had the overhaul shop we used to do a lot of lapping. Had a lapping plate in the engine shop and a flat piece of thick glass as a lapping plate in the accessory shop. Saturate some 800 wet-or-dry sandpaper with light oil and then swirl the cover plate across the surface a few times - then you look at the cover plate and see if it is distorted. If it is then continue lapping.

Another problem with gaskets is too much torque and you crush the gasket. The motivation is to tighten harder if you have a leak. But look at the picture below. More torque on the bolts and you just distort the surfaces more. The correct amount of bolt torque on a gasketed joint is set by the stress needed in the gasket material to effect a seal and sufficient torque to provide equal pressure across the gasket. Tightening should be done in stages to compress the gasket equally.

Lapping Technique:

The cover plate is swirled against the glass and paper. The wet sandpaper will stick to the glass so just pick up the plate and place it onto the sandpaper and lightly swirl the plate. You can lift the plate off and look at the surface and see the contact area. This will tell you if the plate is flat or not. At this stage you are just checking for flatness. If the paper is not removing metal from the entire surface then continue -- hold the plate with light but even pressure - do not force the plate onto the paper as you will press harder on one side than the other.

Here is a video - a couple of comments - outside of a flat-plate (lapping plate) the next best surface to lap against is glass. Do not glue the sandpaper down to the surface - what a mess. Light oil will hold it down. wet both sides of the sandpaper. I have never lapped with water - always light oil but I suppose water would also work.

videos on lapping

Two essential tools missing from almost every aircraft repair shop - an arbor press and a lapping plate!

Flight Training and Cat Health

2011-06-05T09:34:00.000-07:00

The pattern of Google searches for "flight training" correlates closely with the volume of searches for "cat health." If Google can be believed, the interest in flight training and cats closely agrees over a period of many years.

Here is a comparison with other search terms:

0.9703cat health
0.9700cat care
0.9694listing
0.9686feline
0.9685telephone prefix
0.9679telephone codes
0.9673message boards
0.9668indian baby names
0.9666flight schools
0.9666telephone area

Now it gets even more interesting. Here is the correlation with "Cessna" and "feline"

Not as close but still remarkably similar. Here are the top 10 correlations with "Cessna"

Correlated with cessna

Feline #1 and cat care #7

The data suggests that the interest in flight training, Cessna, and cats changes at the same time. Does this strong correlation hold true for other aircraft manufacturers? How about Piper, Beechcraft, or Boeing? No correlation found with cats or feline. Here is the Boeing data:

Correlated with boeing

0.7765boeing aircraft
0.7677boeing airplane
0.7485boeing planes
0.7431boeing jet
0.7136boeing airplanes
0.6950plaza hotel
0.6914arinc
0.6891real estate brokers
0.6887estate brokers
0.6870shutters

Pretty much what one would expect. So what does this all mean? I have no idea, just thought it is interesting.

Aircraft Engine Bearing Analysis - Reading the Bearings

2011-05-07T09:09:00.000-07:00

These bearings are from Lycoming and Continental aircraft piston engines. In no instance is the bearing itself defective. Bearings are victims of abuse and neglect.

Titanium Limitations in Aircraft Repair

2011-03-06T08:10:00.000-08:00

Titanium is entering main-stream usage. The Boeing 787 is 18% by weight titanium. To maintain aircraft that contain titanium it helps to know the material's quirks and limitations...and titanium has a few big ones... such as titanium's fire hazard; Titanium catches fire before it melts - unusual for a metal.

Fire Hazard:
"There have been over 140 known instances of titanium fires in aircraft turbine engines in flight and in ground tests" 1.

Fire damage to titanium and titanium alloys becomes critical above 1000 degrees F due to the absorption of oxygen and nitrogen from the air which causes surface hardening to a point of brittleness. An overtemperatured condition is indicated by the formation of an oxide coating and can be easily detected by a light green to white color. If this indication is apparent following fire damage to titanium aircraft parts, the affected parts will be removed and replaced with serviceable parts. T.O. 1-1A-9 page 5-6

"The application of titanium in the engine design should be directed primarily to minimizing the probability of uncontained titanium fires, i.e., fires that penetrate the engine casing" 1.

1. FAA AC33.4 "Design Considerations Concerning the Use of Titanium in Aircraft Turbine Engines.

Hydrogen Embrittlement
Hydrogen-embrittlement is a major problem with titanium and titanium alloys. Hydrogen is readily absorbed from pickling, cleaning and scale removal solution at room temperature and from the atmosphere at elevated temperatures. Hydrogen embrittlement in the basically pure and alpha alloys is evident by a reduction in ductility and a slight increase in strength. This is associated with a decrease in impact strength at temperatures below 200 degrees F. and a shift in the temperature range where the change form ductile to brittle occurs.

With alpha-beta alloys, embrittlement is found at slow speeds of testing and under constant or "sustained" loads as demonstrated by tests on notched specimens. This type of embrittlement, which is similar to the embrittlement of steel, only becomes evident above a certain strength level. Solution heat treating and aging the alpha-beta alloys to high strength levels increases sensitivity to hydrogen embrittlement.

Cadmium Plate Caution

Cadmium plated self-locking nuts shall not be used in contact with titanium and titanium alloy bolts, screws, or studs in application where the operating temperature exceed 450 degrees F. Cadmium plated clamps, fixtures, and structures per Aeronautical-Design-Standard-ADS-13F-HDBK. Note, when considering localized cadmium embrittlement of titanium, consider that friction can sometimes cause this heating effect.

Boeing-Design-Manual-BDM-1054 states "The use of cadmium plated titanium components is not allowed. Cadmium plated components which come in contact with titanium are not allowed, except for hydraulic systems where cadmium plated steel fittings may be coupled to titanium fittings and cadmium plated steel or titanium nuts on titanium or steel bolts. MIL-S-5002 prohibits all contact between titanium and cadmium on military programs." Cadmium plated clamps, fixtures, and\ jigs should not be used for the fabrication or assembly of titanium components or structures. Cadmium plated self-locking nuts shall not be used in contact with titanium and titanium alloy bolts, screws or studs. MIL-HDBK-1599A.

Silver Plate Caution
Silver-plated self-locking nuts shall not be used in contact with titanium and titanium alloy bolts, screws, or studs in application where the operating temperatures exceed 600 degrees F. Per MIL-STD-1515A "Fastener Systems for Aerospace Applications" Silver brazing of titanium parts should be avoided for elevated temperature applications." ADS-13F-HDBK at temperatures exceeding 230°C (450°F). The warning on cadmium and silver is most likely because it was found that when cadmium or silver plated fasteners were pressed or smeared into the titanium surface at or near the yield of titanium that embrittlement of the titanium and cracking resulted. This became known as cadmium-embrittlement or Solid-Metal-Embrittlement (SME). Any barrier that prevents direct contact (such as a dry film lubricant) can prevent cadmium embrittlement. In most applications, the likelihood of SME is quite low or non-existant since the cadmium must be smeared into the surface while titanium is in tension well above 50% of its yield strength.

Skydrol Caution
Titanium can be embrittled by accumulations of Skydrol-hydraulic-fluid (BMS3-11) at temperatures above 270 degrees F. Per Boeing-Design-Manual-BDM-1054.

Alcohol Caution
Titanium can be embrittled by methyl alcohol and anhydrous ethyl alcohol at room temperature. Per Boeing Design Manual BDM-1054.

Solder Caution
Titanium can be embrittled by silver, zinc, lead and lead alloys at elevated temperatures. Per Boeing Design Manual BDM-1054.

High Temperature Caution
Titanium should not be used at temperatures above 1050 degrees F 565.6 C) as it has an unusually high attraction for carbon, oxygen, nitrogen, and hydrogen above this temperature. This makes the titanium brittle. Working with titanium requiring the application of heat in excess of 800 degrees F., must be performed in a closely controlled atmosphere. The absorption of small amounts of oxygen or nitrogen makes vast changes in the mechanical properties. In gaseous oxygen, a partial pressure of about 50 psi is sufficient to ignite a fresh titanium surface over the temperature range from -250 degrees F to room temperature or higher.

Salt Caution
Titanium is susceptible to stress-corrosion-cracking by sodium chloride or chloride solutions at elevated temperatures. If you are using titanium parts above 450 degrees F (232.2 C), then use a nonchlorinated
solvent and avoid leaving fingerprints.

"An American turbine engine manufacturer recently published a service letter alerting operators that wrapping stainless steel tube assemblies with a chloride-based material, such as neoprene tubing and fibreglass tape to prevent chafing, has resulted in premature tube failure. A chloride-based material breaks down from the presence of high engine temperatures and attracts moisture, resulting in the formation of salts which are highly corrosive to stainless steel tubes. After a period of time, stress cracking develops resulting in failure of the tubes. Additional investigation along the same lines by a foreign engine manufacturer revealed that titanium is also affected by the chemical reaction between chloride-based materials when operating in temperatures in excess of 150 degrees C (302 degrees F).

A related problem is the use of chloride-based packaging material, such as PVC sheeting (plasticized polyvinyl chloride) as a packaging material. This can result in chloride-based residue being left on the component, possibly leading to the sort of failure described above.

In summary, operators are reminded to follow the engine manufacturer's publications in installing stainless steel engine air, oil and fuel tubes and warned against using chloride-based materials on any stainless steel or titanium components, whether installed on the engine or held in storage. " AAC 1-13 Australian-Government-Civil-Aviation-Authority

Mercury Caution
Under certain conditions when in contact with cadmium, silver, mercury, or their compounds, titanium may become brittle. Refer to MIL-S-5002 and MIL-STD-1568 for restrictions concerning applications with titanium when in contact with these metals or their compounds. Silver will cause cracking in many titanium alloys at temperatures above 650 degrees F.

Liquid Oxygen Caution
The use of titanium in contact with liquid oxygen should be avoided since the presence of a fresh surface, caused by cracking or abrasion, may initiate a violent reaction. Per Boeing Design Manual BDM-1054

Wear and Galling Caution
Titanium-galls very easy. It has been described as a "gummy" metal, strong but soft. Titanium threaded fasteners may require anti-seize. The loss of Sikorsky S-92A ship number CHI91 due to galling of titanium studs is an example of how galling is a serious concern. Conversion coatings, such as Tiodize can be applied to titanium fasteners to prevent galling. For example the Titanium interference fit bolts in the F-14 wings would gall if driven into the hole bare. Tiodize coating is used to prevent such galling. Bare titanium should not be used for components having sliding surfaces. Pined joints subject to rotation, vibration, or repeated loads must be bushed with unplated aluminum-nickel-bronze or CRES bushings.

Crevise Corrosion Caution
Titanium is susceptable to crevise-corrosion in chloride (salt) solutions at elevated temperatures. Different heat treatments and alloys vary. "Care should be taken to ensure that cleaning fluids and other chemicals are not used on titanium assemblies where entrapment can occur. Substances which are known to be contaminants and which can produce stress corrosion cracking at various temperatures include hydrochloric acid, trichlorethylene, carbon tetrachloride, chlorinated cutting oils, all chlorides, freons, and methyl alcohol." ADS-13F-HDBK

Galvanic Corrosion Caution
Titanium is similar to Monel (nickel-copper alloy) and stainless steel and galvanic reactions generally will not occur when coupled with these materials. Less noble materials, such as aluminum, carbon steel, and magnesium alloys may suffer galvanic corrosion when coupled to titanium.

Welding Caution
Titanium welding must be done in an inert atmosphere. Cracked titanium bicycle frames are a good example of how lax attention to welding details results in fatigue cracks years down the road. Here is a write-up from a fatigue failure of a titanium duct on a Lockheed Tristar:

Although welding of commercially pure titanium normally results in a slight local hardness increase, a well executed weld should only produce an increase in the range 10 to 25 HV. The weld at the duct fracture location exhibited a much greater hardness increase (45 HV) and would therefore be expected to have had reduced ductility, impairing the fatigue characteristics of the duct. A difference greater than 30 HV compared with the parent material with an associated loss of ductility can indicate that gas contamination has occurred, leading to weld embrittlement. Gas contamination and embrittlement occurs when the weld pool is not sufficiently shielded from atmospheric gases such as oxygen, nitrogen and hydrogen. The blue/purple tint to the weld area adjacent to the fracture is evidence of elevated temperature oxidation... AAIB Bulletin No: 6/99 Ref: EW/C98/9/5 Category: 1.1

This report seems to imply that a local hardness test close to the weld may be a good test for excessive embrittlement.

When titanium is heated to 500 degrees C. (930F), it absorbs oxygen, hydrogen, nitrogen, and carbon. These atoms enter the titanium and make it brittle. Evidence of titanium weld contamination is readily apparent as a discoloration of the weld surface. This discoloration is caused by oxidation and starts at about 900 degrees F. Heating to temperatures above 1000 degrees F. under oxidizing conditions results in severe surface oxidation and brittleness.

General Welding Principles:

Not every good welder can weld titanium - requires discipline.
Cleanliness. You do need to be manic about cleanliness. Solvents must be very fresh and always stored in sealed containers. Purge gasses need to be pure. Avoid rubber or plastic hoses in handling the gasses. The permeability is too high and you will pick up oxygen and moisture. Use Lint-free gloves after cleaning so as to avoid contaminating the surface with perspiration.
Protect the backside. Wherever the titanium is heated, brittle alpha-case can form.
The presence of blue or white oxide is an indicator that contamination has occurred.
A bright silver color (mercury color) is desired.

High quality industrial and aerospace welding of titanium is done in a hermetic welding chamber which maintaines the atmosphere of Argon with less than 20-ppm O2 and 20-ppm moisture.

Hydrogen Migration to Weld at Elevated Temperature Caution
Metallurgical examination of the duct fracture surfaces showed that it had failed due to cracking from multiple origins on the duct inner surface, adjacent to the weld. Hydride formations were present and the metallurgical report concluded that the failure was similar to that described in Boeing Service Bulletin 747-36A2074...This states "At duct operating temperatures of 300 to 350 degrees Fahrenheit, hydrogen in the titanium duct material tends to migrate towards areas of high stress, and then during cooling, hydrides form. These hydrides have an embrittling effect on the duct material and may contribute to crack initiation...Studies indicate that stress relieving the ducts eliminates the residual stress and local stress concentrations which stops the migration of hydrogen to the circumferential welds." Airframe cycles on duct: 14,698. Ref. AAIB EW/A92/6/1 Boeing 747-283B, G-VOYG

Monel Safety Wire - When to Use

2011-03-06T07:23:00.000-08:00

Stainless Steel Safety Wire on left - Monel Safety Wire on right
Hard to tell apart

Monel resembles stainless steel but is an alloy of Nickel and Copper. Probably the easiest method of telling stainless steel from Monel is the spark test. Stainless steel will create sparks when placed against a grinding wheel. Monel (and most nickel alloys) are "non-sparking"

Monel wire is used by deep sea fisherman as a trolling, seizing and baiting wire because it can be rolled and bent many times without breaking - unlike stainless wire. And its extra softness makes it easier to use. Monel safety wire was used exclusively on the Titan II rocket engine (picture below is from a Titan II). Given that the operational life was 3 1/2 minutes for this first stage engine one wonders why safety wire was even needed.

Monel Safety Wire used on the Titan Rocket Engine

Monel is used in potentially explosive-atmospheres where the materials must not be capable of sparking. An example is using monel safety wire in an aircraft fuel tank instead of stainless steel safety wire. Reference Boeing-767 AD-2006-08-04. Although, stainless steel has lower potential to spark from friction than a tool steel, Monel is even better.

Long gap between safety wire ends

In the picture above, the long unsupported gap between safety wire ends and engine vibration may cause the wire to resonate and fatigue fail. Monel safety wire, with its better toughness and fatigue strength, is a better choice than standard safety wire in this application. In this picture the stainless steel safety-wire-fatigue-failure caused a forced landing Airbus-A330-301. 1. The V-band clamp unwound when the safety wire broke and this resulted in 600 degree C. hot air to enter the engine compartment setting off the fire detection system.
Ref: AAIU-Synoptic-Report-No-2006-006

Identification:
Nitric-acid turns metal blue-green. Steel rod rubbed in solution will turn copper colored. Non-magnetic - magnet will not stick. Monel is magnetic whereas "K" monel is non magnetic Stainless Steel and Monel safety wire look almost identical. If you take a piece of wire and hold it to a grinding wheel, the stainless steel will spark and the monel won't.

Advantages of Monel:

Monel is used in high-temperature areas such as on the exhaust. Monel 400 melts at 2,600 degrees F.
Monel is used for locations where you don't want a spark, such as inside fuel tanks.
Monel resists breakage when bent or vibrated much better than stainless.
Monel is an excellent general purpose wire and better then stainless steel as it will bend more without breaking. This make Monel re-usable.
Monel is also a little bit softer so is easier to work with your hands.

Aircraft Control Cable - What is it?

2011-03-05T08:13:00.000-08:00

The standard for fight-critical aircraft control-cable is MIL-DTL 83420. It is estimated (Defense Daily Network July 27, 2005) that less than 2% of "aircraft control cable sold in the world today meets MIL-DTL-83420. Most of it is what you would find in your local hardware store. Tests performed on non-MIL-DTL-83420 cable concluded that the fatigue strength requirements were rarely met. If your log book entry or sales receipt uses the term "aircraft control cable" then you might be implying that the cable is MIL-DTL-83420 when it is not.

There are two easy identification methods that may help you identify aircraft control cable:

All MIL-DTL-83420 contains a two-color tracer filament emended within the cable that identifies the manufacturer,
All MIL-DTL-83420 cable sold on a shipping real must contain the identification number of the manufacturing reel. (All MIL-DTL-83420 cable is lubricated with a corrosion inhibitor.)

Aircraft Control Cable Wear

Discussion:

Aircraft devices are designed based on:

Strength
Endurance

Often we focus only on the strength aspect. "How strong is it? or How many "G's"? One should also ask "For how long? This is called "endurance." Fatigue strength gives us endurance. The principle difference between aircraft and non-aircraft control cable is endurance (fatigue strength). How many times will the cable bend over the pulley before it starts to break (frays)? Fatigue strength is measured in number of cycles at a given load.

One may think incorrectly that fatigue strength is not so important for a lightly loaded aircraft control cable. An interesting example of fatigue strength importance is on the Eagle-Aircraft where the control cables were fraying between 400 and 900 hours in service. In models X-TS150 and 150B, Australian-Airworthiness-Directive-AC/XT-S/2 and CASA #0008 was issued along with Service Bulletins from the aircraft manufacturer to inspect cables for fraying at the pulley. In this instance the fraying was attributed to the small size of the cable-pulley. Although not mentioned in the report, cable fatigue strength is also a factor in cable fatigue failure (fraying).

Another possible cause is the use of stainless steel cable instead of galvanized steel. Stainless steel has high friction and the individual wires can gall as they rub against one another. Galvanizing acts as a lubricant and keeps the steel cable from wearing. Consequently, wear rates on stainless steel cable used where the wires may move - such as rounding a pulley - are far greater with stainless steel. More frequent inspections are required. As of 2004, Boeing uses practically no stainless steel cables. They use the Tin over Zinc variety of carbon steel cable in their primary flight control cables.

A review of Malfunction and Defect reports from several countries seems to show that premature cable fraying is not an isolated event. A fatal Twin-Otter-crash because of worn stainless steel elevator cables in Tahiti prompted BAE, Transports Canada and the European-Agency-for-Air-Safety to ask owners of these aircraft to inspect elevator cables. This is not a new problem. In our relatively lightly loaded control systems, cable fatigue strength and wear rate might be more important than ultimate strength.

Stainless Steel Stress Corrosion Cracking - Primer for Aircraft Mechanics

2011-03-05T08:01:00.000-08:00

Stress Corrosion Cracking

What does an indoor swimming-pool-roof-collapse in Switzerland that kills 14 people, an aileron failure in a Bonanza, and rock-climbing-bolts that break when slightly tapped have in common?

Failures occurred in parts made from 300 series (austenitic) stainless steel. The most commonly used grade.
Parts were exposed to salts and chlorides
Parts were stressed in tension.
Engineering standards and tests at the time said it couldn't happen.

Seven aircraft have lost flight control because the stainless control cable terminals cracked due to corrosion. The turnbuckle terminals (part number MS21250 or AN669) are made from 303 stainless - a common grade of stainless. This type of terminal is used on most general aviation aircraft and helicopters. Piper reportedly manufactured 51,600 airplanes containing these terminals. One Navy aircraft suffered a failure. There is also a long history of turnbuckle breakage in sailboats.

When chloride-salts get into crevices where there is a lack of oxygen, pits form in the stainless and the part eventually breaks from the inside out. This is called Chloride stress-corrosion cracking. Since the corrosion forms pits inside crevices, the part may look perfectly good from the outside. In the case of the turnbuckle terminals, general corrosion pits were found on the surface of "most" of the broken terminals. Also, in the AN669 series, the safety wire wrapped over the terminal hid the corrosion pits.

In the case of the rock climbing bolts, they looked fine until lightly tapped and broke flush with the rock face. Pretty scary if you are dangling from one of those bolts. In the roof collapse, the stainless hangars were above the ceiling panels hidden from view. 300 series stainless is now banned in the European Union, Switzerland, and Australia for use in indoor swimming pools when used in safety critical applications. It is still being used on aircraft flight controls!

Stainless steel (especially the common 300 series) does not like chlorides. Chlorides are found in salt water, road salt, and some cleaning solutions such as trichloroethane, and methylene chloride. (trichloroethane is often found in the cleaner portion of dye penetrate cleaners that are used in the aircraft industry to find cracks.) Some insulation material contain chlorides. The worst corrosion combination for stainless steel is low-oxygen and high chlorides as might be found in crevices.

Salt-deposits are hygroscopic, they absorb moisture from the air. When the relative humidity is over 50%, the surface becomes wet and corrosion starts. Wash off any salts that may have been deposited on you're equipment.

Failure Characteristics:

SCC failures can occur rapidly or very slowly. Inspections or replacement based on time-in-service may not be a useful criteria.
SCC failures are rapid, complete break of the part. There is no tell-tale bending or sagging.
Visual inspection has not been helpful in identifying suspect parts before failure.

Best Maintenance Practices:

It appears that the best maintenance practice is to keep the parts clean so that chlorides don't concentrate on the surfaces.

Best Engineering Practices:

Use better stainless grades, such as the Superaustenitic or Duplex grades.
Use shot-peening to improve the SCC resistance. Shot peening is a proven method of improving the SCC resistance in austenitic stainless steel parts.

Other areas to be concerned about:

Load bearing stainless parts exposed to chlorides where the failure could result in a safety hazard. Some examples might include:
Aircraft structural parts, such as bolts, turnbuckles, etc. where aircraft are based or operated next to the ocean
Swimming pool ladders and bolts used on swimming pool slides and diving boards.
Bolts used on trailer hitches on vehicles next to the ocean or where salt is used on the roads to melt ice.
Bolts used on road signs next to the ocean or where salt is used on the roads to melt ice.

Aircraft Control Cable - Stainless or Galvanized?

2011-03-05T07:29:00.000-08:00

Aircraft Control Cable
Galvanized on top and Stainless on bottom

"A general service history has shown the use of stainless steel cables in aircraft control systems results in premature wear and has been a factor in minor incidences as well as catastrophic failure. A current trend is underway in the aviation industry to move away from the use of stainless steel cable for primary flight control applications, except where marine operations are performed." FAA Special Airworthiness Information Bulletin CE-12-01 Dated October 24, 2011.

Frayed stainless steel control cable is suspect in a Twin-Otter crash killing 14 passengers in Tahiti in August of 2007.

General aviation aircraft generally use control cable made from either stainless steel or galvanized steel. Each type has its advantages and disadvantages. Generally;

Galvanized rope is stronger.
Galvanized rope has greater fatigue strength.
Galvanized rope has less wear.
Galvanized rope is easier to inspect for corrosion damage.

Per FAA CE-12-01:
Stainless steel is more corrosion resistant.
Stainless steel has considerably less service life due to high wear.
Stainless steel cannot be inspected for corrosion damage.
Stainless steel is stiffer and has lower bending fatigue resistance - important in flight control systems
Stainless steel has a higher friction coefficient that results in increased wear every time the cable is flexed.
Stainless steel becomes more stiff, leading to increased abrasion wear in the inside as well as the outside of the cable

The poor wear resistance of stainless steel rope has resulted in aircraft control problems. More frequent inspections are required. For more information on this subject reference: Special-Airworthiness-Information-Bulletin:-SAIB CE-01-30, July 11, 2001. For 172S airplanes see FAA SAIB: CE-11-3. also CE-11-36, Piper Service Bulletin 1048.

There are several reasons why stainless wears more than galvanized steel when used on flight controls:
The bending of a wire rope causes the individual wire stands to not only bend but to rub against one another. Galvanizing is a natural lubricant. For example, galvanized threads have a lower friction (K) factor then plain steel. The individual wires can easily move about with very little friction and wear. Stainless steel on the other hand has high friction and has a reputation for seizing and galling when rubbed together. Every time the wire rope is flexed, the stainless wires rub together. High friction creates high wear.

There are several methods of reducing wear and increasing fatigue resistance in a wire rope. Wear resistance can be increased by changing how the wire stands are wound. In the picture above the individual wires are horizontal (parallel to the axis of the rope). This is called "right regular lay" and is the standard lay. Another method of winding the wire stands is so that they form an angle to the axis of the rope. This is called "lang lay". Lang lay increases fatigue strength and abrasion resistance without any decrease in ultimate strength. Another method of changing the wire characteristics of fatigue strength,. abrasion resistance, and flexibility is to use wires of different diameters. For example, Douglas-Specification DMS2192 calls for a Warrington Seal (IWRC) construction. This type of wire rope has larger wires on the outside and and smaller wires on the inside.

There are other wire rope designs that the engineer can call for to optimize specific performance goals. This is why when we replace wire rope we should make sure that the replacement meets the original manufacturer's specifications.

Galvanized Vs Stainless - who uses what?
As of 2004, Boeing uses practically no stainless steel cables. They use the Tin over Zinc variety of carbon steel cable in their primary flight control cables.

Aircraft Wire Inspection

2011-03-05T07:23:00.000-08:00

Rats!

ASTM Standard F 2696 08 "Standard Practice for Inspection of Airplane Electrical Wiring Systems" is an excellent resource for developing a wire inspection system. Some comments I have are below:

"The principal technique for inspecting aviation and spacecraft-wiring components used to date remains visual inspection. These inspections are unable to detect all extant flaws and are subject to discrepancies and errors. Moreover, these tests are intrusive, since brittle wire bundles are frequently moved to access more remote wiring components and can result in further damage to already cracked insulation."

Source: NONDESTRUCTIVE EVALUATION OF AROMATIC-POLYIMIDE-INSULATED AIRCRAFT AND SPACECRAFT WIRING. E.J. Tucholski, Phusics Department, U.S. Naval Academy, Annapolis, MD

It should not be expected that the mechanic will be able to detect all aircraft wiring faults through visual inspection. It's the responsibility of the FAA and aircraft engineers to develope suitable inspection tools or apply age control limitations to aircraft wiring. The problem of "aging aircraft wiring" is an age problem and not a maintenance problem.

"When a failure could have catastrophic results, it is not appropriate to rely on maintenance and inspection intervention to prevent the failure from occurring if a practicable design alternative could eliminate the catastrophic effects of the failure mode." Quote from NATIONAL TRANSPORTATION SAFETY BOARD Public Meeting of December 10, 2002 Abstract of Aviation Accident Report Alaska-Airlines-Flight-261, MD-83, N963AS NTSB/AAR-02/01

Wire Support using red RTV

Inspection tips:
High voltage and low voltage fuel sensor wires should not be mixed in the same bundle. Damage to the wire bundle could allow high voltage to enter the fuel tank. Recommendation resulting from TWA-Fligh- 800 747 fuel-tank-explosion

No wiring is routed in proximity to oxygen, fuel, and hydraulic lines or critical flight control cables. ASTM F2696-08 reads in part: "where practical, route electrical wires and cables above fluid lines and provide a 6 inch (15 cm) separation from any flammable liquid, fuel, or oxygen line, fuel tank wall, or other low voltage wiring that enters a fuel tank and requires electrical isolation to prevent an ignition hazard. Where this 6 inch cannot be maintained then wiring should be closely clamped and rigidly supported to avoid contact even assuming a broken wire or missing clamp.

Protect the wire from contamination by fluids (including corrosion inhibiting compounds), flammable lint, metal shavings, or other debris. Hard materials can work their way into the wire bundle and with vibration penetrate the insulation causing electrical shorts. Fluids can soften or crack the insulation. Water and dirt become slightly conductive and lead to arc-failures.

Unless advised otherwise in the maintance manual, do not mix wire insulation types in the same bundle because insulation of different hardness may create chafing damage in vibration areas.

Wires and cables are supported by suitable clamps, grommets, or other devices at intervals of not more than 24 inches (61 cm).

Wires must be grouped, routed and spaced so that damage to essential circuits will be minimized if there are faults in heavy current-carrying cables. The objective is to minimize the impact of the failure of a heavy current-carrying cable on any essential system wiring.

Protect the wires from moisture and high temperatures; these cause wire insulation to age and crack.

The minimum radii for bends in wire groups or bundles shall not be less than ten times the outside diameter of their largest wire, except at the terminal strips where wires break out at terminations or reverse direction in a bundle.

Due to cold flow phenomena of teflon insulataion used in MIL-W-22759 wire, it is advised NOT to route teflon insulated wires over sharp edges and tight turns, or apply tight tie wraps to cable assemblies. Cold flow or creep is the slow movement of the insulation when under a steady-state stress. The old practice of using soft Koroseal-lacing (rubber lacing) to tie wire bundles together is much less damaging then using hard plastic tie-raps.

If you are developing a wire inspection standard for your business I would suggest that you review ASTM Specification F2799-9 "Standard Practice for Maintenance of Airplane Electrical Wiring Systems." and ASTM Specification F2799-8 "Standard Practice for Inspection of Airplane Electrical Wiring Systems" These are excellent documents for any shop and are well worth the price.

Clean and Dry Torque

2011-03-05T06:57:00.000-08:00

"Clean and Dry" Problem for the Mechanic - Problem for NASA

Is this stud "clean and dry"?
How should it be cleaned?
MIL-HDBK-60 offers guidance

Bolts and studs are often plated with Cadmium, Zinc, and other coatings that have published "K" (friction) factors. Torquing the bolt using a "clean-and-dry" specification produces a reasonably accurate amount of tension based on these 'K" factors.

With new bolts and studs there are no cleaning issues for the mechanic as the bolt or stud is received in the "clean and dry" condition. What about old used bolts and studs? They are not received in the "clean and dry" condition. In what manner should the mechanic clean the surfaces? Should a wire brush be used to clean old thread locking compound from the threads?

Plating may be damaged or worn; threads might be damaged; rust, paint, and adhesives might be stuck in the threads. Cleaning often involves whatever is handy, such as a wire wheel or wire brush. Whatever plating is left in the threads might be worn off during cleaning. The cleaning compound might have a big impact on thread friction. 1

Threads create 50% of the friction resisting torque, the bearing surfaces create the other 50%. Does "clean and dry" apply to only the bolt threads? Or does it apply also to the bearing surfaces?

A "clean and dry" torque specification for used bolts and studs without specific cleaning and inspecting directions is deficient. The "K" factor on some old used bolt and stud might be almost anything and vary from bolt to bolt. Not only engineers but mechanics should be aware of this limitation.

NASA found this out during testing of a model wing in their 8' Transonic-pressure-tunnel when the flap peeled away from the wing, broke free, and proceeded down the tunnel. Engineers had specified a "clean and dry" torque. This was quite impossible as the assembly directions also specified that a liquid thread locking compound be applied to the bolt during assembly.

The particular bolt was often removed and reinstalled during testing. What the technicians at NASA did was re-apply thread locking compound to the bolt each time it was installed. NASA lost the model wing when the bolt backed-out. NASA's "Lessons Learned" document states quite simply: It is impossible to predict torque value on screws after repeated applications of "a thread-locking compound".

Why would a "clean and dry" specification be used?

Clean and dry threads and bearing surfaces have greater friction than lubricated surfaces. Friction helps prevent the bolt or nut from loosening and backing-off. Clean and dry, uses friction to our advantage and can be an aid keeping fasteners tight.

MIL-HDBK-60-THREADED-FASTENERS - TIGHTENING TO PROPER TENSION offers this description of "dry": "So-called "dry" threads refer to threads where no lubricant is applied. Some residual machine oil is assumed. If all lubricant is removed by solvent, coefficient of friction is inconsistent and often very high unless a plating or other film is acting as a lubricant. Severe galling may also result from lubricant-free surface conditions."

It appears from the above quotation that a proper torque condition statement is "dry threads" rather than stating "clean and dry".

1. "Failure-of-bolts-in-helicopter main rotor drive plate assembly due to improper application of lubricant" by N. Eliaz, G. Gheorghiu, H. Sheinkopf, O. Levi, G. Shemesh, A. Mordecai, H. Artzi, Published in Engineering Failure Analysis #10, 443-451.

Double Flare Tubing

2011-03-04T13:03:00.000-08:00

Aircraft Standard MS33583 Double Flare

"A double-flare is used on soft aluminum tubing 3/8 inch outside diameter and under, and a single-flare on all other tubing." AC43.131B "FAA Acceptable Methods, Techniques and Practices"

As system pressure increases, tubing joints must be designed to withstand these pressures. 5052-0 is soft tubing and the flare is not strong enough to handle higher pressures. Double flaring reduces cutting of flare by overtightening and failure of tube assembly under operating pressure. A double flare is stronger in fatigue. Except in emergencies, there is no acceptable reason to use a single instead of a double flare where appropriate. That said, in practice, very few flares made by mechanics are double flares, however, this does not make the practice acceptable nor desirable.

A leak-free connection is not the sole measure of an acceptable flare. The proper radius at "B", and no nicks or other damge is required for a durable connection that won't break later in service. This is why using a 45 degree flaring tool and then mashing the flare to a 37 degree with the "B" nut may produce a leak-free connection for the moment but does not produce a safe, durable connection.

Construction Standards per MS33583

Double Flare Radius. Screenshot from Mechanic's Toolbox Software

B Radius

0 1 2
Tube Size Outside Diameter, Inch B Radius, Inch
1/8 0.032
3/16 0.032
1/4 0.032
5/16 0.032
3/8 0.046
1/2 0.062
5/8 0.062
3/4 0.062
1 0.093
1-1/4 0.093
1-1/2 0.109
1-3/4 0.109
2 0.109
2-1/2 0.109
3 0.109

Measuring B radius

Measuring B Radius

Continental Cylinder Stud Design

2011-03-04T07:51:00.000-08:00

Continental just makes a better cylinder ...I have no vested interest in Continental and I don't sell their cylinders but I do appreciate good engineering and design...attention to details that I don't see in other "PMA" cylinders.

PMA Cylinder Workmanship

Notice the stud isn't even straight and the design puts the maximum stress at the surface where it is pulling metal already.

Continental Workmanship

Now look at the same stud on a Continental cylinder. Chamfered hole. Notice how the first engaged thread occurs below the surface. Stud is "waisted" - that is a good thing!. Waisting is the reduced diameter in the unthreaded portion of the stud. This diameter is now the same as the root diameter of the thread making the stress evenly distributed throughout the stud.

Wasted Cylinder Hold-Down Stud

Here is a picture of a waisted cylinder hold-down stud on a Continental. The reduced shank diameter also allows the stud to store more energy as it will stretch more than a non-wasted stud for the same amount of applied tightening torque. This increases fatigue strength and helps prevent joint loosening.

Waisted Cylinder Hold-Down Studs on Continental IO520 engine

Graphite Lubricants in Aircraft- The Corrosion Potential

2011-03-03T08:27:00.000-08:00

Graphite Antiseize

"...shown conclusively that graphite in a resin-bonded solid film lubricant is deleterious from the point of view of corrosion protection provided by the lubricant... To use graphite is to invite corrosion difficulties in the presence of moisture." Rock-Island-Arsenal-Lab, Technical Report, Dry-Lubricants and Corrosion, Prepared for Presentation at the Annual Meeting of the Society of Automotive Engineers, Detroit, Michigan 14-18 January 1963. Francis S. Meade and George P. Murphy, Jr.

Aircraft Spark Plug Graphite Antiseize

Graphite and water has been used and recommended for aircraft spark plugs for over 50 years. Other anti-seize that contains metallic particles have been avoided because of the chance that the anti-seize particles may get into the combustion chamber, create a hot spot, and cause destructive pre-ignition.

Years ago mica-anti-seize was commonly used for spark plug anti-seize and it probably has better properties than graphite and water. However, there is a problem in changing anti-seize types; you change the torque tension relationship. Published torque values in aircraft spark plugs are based on using the manufacturer's recommended anti-seize. Use a different type of anti-seize and the required torque to achieve the proper amount of tension will change by an unknown amount.

Cleaning spark-plugs using glass-bead shot may remove the nickel plating from the threads. Now the steel threads are exposed to the water and graphite mixed anti-seize creating a corrosive environment.

When the steel threads rust, their surfaces expands and this causes the spark plug to seize in the threads.

Consideration should be given to using a different type of anti-seize in this special circumstance such as a mica based anti-seize.

Don't Mix your Metals

2011-03-03T07:57:00.000-08:00

Stainless steel hardware (rivets, bolts, screws) installed into an aluminum fitting creates a potential for dissimilar metals (galvanic) corrosion.

Corrosion Pit from galvanic corrosion

Leaking hydraulic line from aircraft brake system caused by galvanic corrosion. Aircraft brake systems are drenched in electrolyte (dirty water) so extra caution is advised when mixing metals. Notice that the gap between the tubing and sleeve can trap moisture into the crevice.

Stainless Steel sleeve on aluminum fitting = galvanic corrosion

"Stainless steel parts are cadmium plated and primed if they are attached to aluminum or alloy steel parts." Boeing Aircraft Aero No. 07 "Design for Corrosion Control"

"Aluminum structure shall be insulated from non-aluminum fasteners" Navy Ships' Technical manual Chapter 075 Fasteners, page 75-59.

"Dangerous corrosion will result if steel, corrosion-resistant steel, Monel, titanium, copper, or iron rivets are used in riveting aluminum structures. Such applications should be confined to extreme emergencies" US Air-Force-Airframe-Repair-Specialists (AFSC 42755), Repair Procedures, page 16.

.Boeing's Big Recall:

News article from King5 News in Renton Washington November 24, 2008 concerning stainless-steel-nutplates not being coated with cadmium. Now, tens of thousands of others lack an important coating of cadmium. That nearly invisible coating is important because it prevents the stainless steel nutplate from reacting with the airplanes aluminum, which can lead to corrosion. Spirit says the untreated nutplates from one of their suppliers got mixed in with treated plates from another were installed by the thousands. Spirit employees are now inside Boeing plants trying to find and replace the bad nutplates on new jets.

Why this Matters

The reason corrosion is such a problem in load bearing structures is that the corrosion pitting provides the perfect nucleation points for fractures to form and propagate from. They must be repaired promptly and properly.

Beech Marks, Fatigue Failure, and High Compression Pistons

2011-02-17T08:12:00.000-08:00

"Designs will fail if subjected to overload...that's just the nature of efficient design--

they might not fail immediately; but some statistically determined time in the future--

they will not fail from overload but from a more insidious process called fatigue."

Lycoming Crankshaft Fillet - Fatigue Failure

Beech Marks are a sign that a crack progressed across the part and failure was due to fatigue. They are shown in the picture at the red arrow. The white arrow shows the crack initiation point.

Fatigue occurs when the metal is subjected to repeated or alternating stresses not exceeding the material's static yield strength. A fatigue failure is a failure due to repeated stress BELOW the material or parts ultimate tensile stress. A part can operate normally and then suddenly fail in fatigue if cyclic stresses are above the fatigue strength of the metal.

Even more interesting is that fatigue strength is a probability based on statistics and not one set value. The actual fatigue strength of a particular part might be less or might be more. There is no way of knowing unless you test it to failure.

So what has this to do with high compression pistons? Anytime you increase engine power above what the engine was designed for you assume that the original design is inefficient -- built stronger than need be and this extra strength caused by design ignorance is just waiting for some smart person to exploit.

But could it be that the original design is competent. That the engineer designed for endurance; the designer knew that statistically some of the parts would be slightly weaker in fatigue so he designed beyond 3 sigma as he had to be sure your crankshaft would not fail; that he designed for an infinite fatigue life by purposely limiting the stress.

So how do you know if the high compression pistons place the fatigue life into the finite part of the fatigue curve without doing the stress analysis? You don't. Will it fail? You don't know. When will it fail? You don't know.

Years ago when I asked a Lycoming engineer what he thought of a popular engine modification he said: 'ask me in 5 years; but of course it might fail in the 6th year." Who knows without the stress analysis.

19th century economist Frederich Bastiat:

"This explains the fatally grievous condition of mankind. Ignorance surrounds its cradle: then its actions are determined by their first consequences, the only ones which, in its first stage, it can see. It is only in the long run that it learns to take account of the others. It has to learn this lesson from two very different masters—experience and foresight. Experience teaches effectually, but brutally. It makes us acquainted with all the effects of an action, by causing us to feel them; and we cannot fail to finish by knowing that fire burns, if we have burned ourselves. For this rough teacher, I should like, if possible, to substitute a more gentle one. I mean Foresight."

Continental IO-520 thrown connecting rod

To summarize: it is better to learn from foresight than experience.

Aircraft Fuel Flow Transducers - Hose Suggestions and Warnings

2011-02-12T08:03:00.000-08:00

Two Rules of Hose Installation that are often Violated when Installing Aircraft Fuel Flow Transducers

1. Fittings should not be used as a bracket.
The fitting should not see any forces (from the overhang weight of the transducer or movement or vibration). A hose should be installed with a slight loop or radius to absorb any movement or thermal expansion or contraction. Some suggest using steel fittings instead of the traditional aluminum AN fittings because they are stronger. If you intend to use the fitting as a bracket then yes, by all means use steel. If you intent to use the fitting as a fitting then steel just adds weight. A copper-based alloy fitting is the ideal fitting to use in aluminum bosses; as a substitute, aluminum.

2. A hose should have a loop or radius. It should not stretch straight from fitting-to-fitting.
A short straight hose is a rigid connection. Any thermal expansion, contraction; movement is transfered to the fitting, thereby violating rule #1 above.

Bulkhead Fitting Installation into Aircraft O-Ring Port

2011-02-10T06:56:00.000-08:00

Aircraft AND10050 Port with Bulkhead Fitting

Low Pressure - Other Than Hydraulic and Pneumatic

1. Assemble AN924 nut onto fitting end and run all the way back to clear fitting groove.
2. Coat male threads and O-ring sparingly with system lubricant.
3. Hold O-ring firmly against the top of the threaded section of the fitting and run nut down until it contacts the O-ring.

Position O-ring and Nut as Shown

4. Turn the fitting into the AND10050 boss and, at the same time, keep the AN929 nut turning with the fitting until the O-ring contacts the boss. The point can be determined by a sudden increase in torque.

Turn the fitting

O-ring contacts the boss

5. Continue to screw fitting into the boss for another 180 degrees. Any further positioning of the fitting must be accomplished by turning the fitting in up to an aditional 270 degrees or by backing out up to 10 degrees. Keep the AN924 nut turning with the fitting to prevent cutting the gasket with the fitting thread.

Fitting installed

6. Tighten locknut lightly
7. Now assemble flared tube to nipple end of fitting.
8. Now tighten lock nut against boss

High Pressure or Hydraulic System Installation

AN6289 Nut

This process uses a different lock nut, a AN6289 nut with a groove for a anti-extrusion device.

Backup Ring (anti-extrusion)

The above description follows AND10064 "Fittings, Installation of Flared Tube, Straight Threaded Connectors" Also reference FAA Special Airworthiness Information Bulletin, SAIB: CE-07-46 dated September 6, 2007 for important installation background and details.

Camshaft Lobe Pitting Evaluation

2011-02-07T07:36:00.000-08:00

I’ve enjoyed reading your articles very much. They are very informative and helpful. I have an additional question regarding cam Spalling. I have a Lycoming IO360 A1A that we removed the cylinders due to a broken ring. While inspecting the cam we noticed minor Spalling on one of the lobes. All the lifters and other lobes look good. The spalling is limited to a single line that runs across the one lobe. The attached picture is not of my actual lobe, but the area circled in RED is representive of the level of damage on my lobe. I’m trying to make an informed decision on to either place the overhauled cylinders back on or major the engine. Can you provide any insight on how long it will be before my cam deteriorates to the point it is no longer airworthy? If this cam will last another 400 hrs I would prefer to leave it alone for now, but if it is only going to last 50 -100 hours I would go ahead and major the engine. Any advice will be very much appreciated.

Camshaft Lobe Pitting

Mike,

The engine manufacturer should be consulted as to the limitations for continued airworthiness.

Assuming the cam follower face is OK? Did you reach in with your hand and rub your fingernail across the surface to detect pitting?

I want you to look at something else on the cam lobe; do you have polishing wear across the entire lobe - end-to-end? Using your exemplar picture notice how the lob surface is shinny from edge to edge. If it has then I would replace the camshaft. The reason I say this is that lobe wear leads to a reduction in power which is an airworthy condition.

Slight pitting does not hinder the proper function of the camshaft but it will progress until it does at an indeterminate rate; start budgeting. Along the way I would use the oil filter inspection technique (originally developed by Lycoming) to detect cam lobe trauma. Do not cut-out the oil filter media but place it into a can with solvent and rinse. Use a toothbrush to lightly scrub between the pleats so any debris is removed from the pleats. Next pour the solvent through a coffee filter and allow to dry. Take a small magnet under the filter paper move all of the magnetic particles from the other debris. If you have enough small metal bits to cover the end of a stick magnet then your lobes and tappets are in a state of active disintegration and the problem needs to be corrected before further operation. Hopefully, you will have none or maybe a stray bit or two indicating that the lobes and tappets have stabilized.

I don't know if this camshaft will last 50, 150, or 400 hours. I would guess that the wear (damage) rate follows roughly an exponential curve. Long duration of little damage and then as the surface starts to pit the damage rate accelerates. Therefore, inspection intervals should be progressively shortened once the onset of pitting is detected.

When to Use a Washer

2011-02-03T08:40:00.000-08:00

Flat Washer with Split Lock

A washer is often used under the nut or bolt, whichever is turned during the tightening operation. When both nut and bolt can be turned, washers are commonly used under both. All washers shall be made from a material which is capable of accepting the peak fastener load without deformation.

A washer can provide multiple functions, the two most important ones are:

1. Spreads the clamping force over a larger area to avoid compressive yielding, and

2. Hard, smooth, consistent material for good preload (clamping) control.

Other functions are to:

1. Prevent galling of the nut face or surface during tightening.

4. Increase energy stored in bolt by using a longer bolt. This helps retain clamping force.

5. Adjusting grip length.

Washer compressive strength must be matched to the bolt/nut clamping force.

WASHER compressive strength MUST be matched to the BOLT/NUT combination! Pictured above is a low-yield strength hardware store washer placed under a propeller bolt. Low-yield strength washers that score/crush in-service under high strength BOLT heads or NUTS, will relieves clamping force, eventually resulting in propeller detachment during operation.

Aircraft Accessory as Designed!

Hydrogen Embrittlement

2011-02-03T08:29:00.000-08:00

652541 nut used on Continental TSIO-520M engine

It is not uncommon to have decorative engine hardware cadmium plated and "oven-baked" during overhaul. Extreme caution is advised. A particular hazard of cadmium-plating high strength steels is the absorption of hydrogen into the base metal. This hazard is countered by baking the parts after plating.

The baking process is critical. Shown above is a hydrogen embrittled nut failure on an aircraft engine. Even though controls were in place, the nuts still failed. Not only is oven temperature and time important, but also the distribution of heat throughout the batch of parts. Oven bake shall occur within four hours of plating and for types II and shall be done before application of supplementary coatings.

Since high-strength steel parts are subject to hydrogen-embrittlement during any plating process they should not be plated unless proper engineering and quality controls have been established and approved. Per T.O.1-1A-9, "All steel parts having a hardness of Rockwell C40 (180,000 PSI) and higher shall be baked at 375 +-25 degrees F. for three hours minimum. SAE-J1648 states: "It may be necessary to provide coatings other than electroplating for fasteners with hardness above 40 HRC"

All steel parts having an ultimate tensile strength of 220,000 PSI or above shall not be plated, unless otherwise specified. When permission is granted, a low embrittlement cadmium plating bath shall be used. Federal-Specifications-QQ-P-416 should be used for cadmium plate requirements. Critical parts should be magnafluxed after plating."

Years ago, the standard was to oven bake for four hours. This was found to be insufficient and the standard changed (2006, but adopted in the 1980's) is that "cadmium-plated parts must be baked at 375 degrees F. for 23 hours, within two hours after plating, to prevent hydrogen embrittlement."

As Cadmium plating is being phased out due to environmental concerns, zinc is often specified as an alternative coating. However, as Lycoming found out, substituting Cadmium for Zinc can lead to disaster. Zinc plating can also lead to hydrogen-assisted cracking. A change from a Cadmium plated crankshaft gear bolt part number STD-2209 to a Zinc-plated-bolt resulted in several aircraft accidents, at least one with multiple fatalities (NTSB IAD02FA091). Several Airworthiness-Directives were issued to remove the Zinc plated bolts and replace them with Cadmium plated ones (AD2002-23-06, AD2002-20-51).

Hydrogen embrittlement and hydrogen-assisted cracking remains difficult to control and predict. There is increasing use of mechanical applied zinc coatings that eliminate the plating process and the resultant hydrogen problem during manufacturer. For further information see the following SAE publication: SAE AMS 2759/9B "Hydrogen-Embrittlement-Relief (Baking) of Steel Parts"

Coarse Thread vs Fine Thread Strength

2011-02-03T08:14:00.000-08:00

Thread Strength Comparison - Fine vs Coarse

Aircraft generally use fine thread fasteners due to their stronger strength. Coarse threads are used when threaded into aluminum or cast iron because the finer threads tend to strip more easily in these materials. The chart above is based on MIL-B-6812E Table II and 125,000 psi UTS.

Aluminum Corrosion Penetration as a Function of pH

2011-01-30T09:36:00.000-08:00

Aluminum Corrosion Penetration as a function of pH

Possibly the best corrosion prevention for aluminum is a neutral pH water wash to eliminate build-up of alkaline salts and then control of pH in the range of 4.0 to 8.5. The protective oxide film that protects aluminum from corrosion is stable and naturally self-renewing. Many cleaners are either alkaline or acidic (citrus based). Soapy water is highly alkaline. At pH values greater than 10, the oxide film starts to dissolve, resulting in rapid corrosion unless controlled by inhibitors.

How to Remove Needle Bearings

2011-01-29T06:25:00.000-08:00

Inexpensive Measuring Stick for Lycoming Pushrods

2011-01-28T18:43:00.000-08:00

$10.00 at Harbor Freight!

Measuring Lycoming Push Rods

One of my customers brought this in today to sort through my box of Lycoming push rods. For $10.00 you can't beat it.

Concord Battery RG Owners Manual

2011-01-19T08:04:00.000-08:00

Concord just released an updated owner/operator manual for their RG series of aircraft batteries.

http://www.concordebattery.com/otherpdf/5-0324-rg-manual.pdf

Slick Magneto Timing Light Flicker

2011-01-15T07:19:00.001-08:00

When timing my slick magnetos I noticed that when the timing light came on and I rotated the prop just a little further the light went out as it should and if I bumped the prop again the light came back on again. What do you think the problem is?

A little dirt or oil on the point surfaces; or a bit of point surface erosion like what is shown in the picture below. A little flicker of the light is probably OK. More and you might need to clean or replace the points.

Inspecting Aircraft Hose

2011-01-07T08:54:00.000-08:00

Comparison of Tube Strength for Common Small Aircraft Tubing

2011-01-07T07:37:00.000-08:00

Design for Strength

Comparative Tube Strength, 6061,5052,2024,6061,3003

Design for Endurance

Since aircraft are vibrating creatures, fatigue strength is also a limiting factor.

Tube Fatigue Strength

Most tubing-failures on aircraft are caused by fatigue. Tubing on aircraft vibrates. How well your tubing endures when subjected to load reversals, impulses, and vibration is called "fatigue strength". Of the three popular tubing types (3003-0, 6061-0, 5052-0), 5052-0 has the best fatigue strength.

Originally aircraft used soft copper-tubing. There is even some of this still around. Although copper was strong enough, it was replaced with aluminum and stainless tube because of the high fatigue failures of copper. For lower pressures, 5052-0 became the tubing of choice because it has the best fatigue strength of any of the non heat-treat aluminum alloys. See "Fatigue Failures of Copper Alloy Fuel", AWB 28-007

--editorial--
Copper tubing on older aircraft should be removed and replaced with 5052-0 before it breaks. There is no warning when copper tubing breaks. One cannot "inspect" it and declare it ok. Another limitation on the use of copper tubing in aircraft engine compartments is that copper strength decreases rapidly with temperature.

Some experimental and light-sport aircraft have hydraulic and fuel lines built with 6061-0 or 3003-0-tubing. Low ultimate strength and low fatigue strength provide a narrow safety margin in dynamic (vibration or impulse) applications. Take extra care in clamping and preventing tube vibration. The aircraft industry's long experience with copper tubing failures proved the importance of fatigue strength. 5052-0 has higher strength and higher fatigue strength at a small price difference. Both 5052-0, 3003 have the same Cold Workability Rating of A (easy to work with). 6061-T6 has a far lower rating of C. "It hasn't failed yet," was the attitude at NASA that essentially led to both of the Space Shuttle disasters; the complacency arises from skirting the line and surviving. But the law of large numbers will eventually get you.

The aluminum hydraulic lines on the Cessna-404 have experienced 5 reported failures due to metal fatigue. For the mechanic, this means that these lines cannot be inspected for fatigue failure. They will not show fatigue stress before failure. A replacement interval is the only method of prevening failure. The old adage that "if it flew in it will fly out" only works until the next failure. See NTSB Safety Board Recommendation A-83-1-2. Metal fluid lines in aircraft subject to vibration have a potential to fail due to metal fatigue. Using the proper alloy tube, combined with good fabrication techniques, and proper clamping, and hard-time replacement interval is the only protection from sudden failure due to metal fatigue.

Diameter Effects

The larger the tubing diameter, the less pressure it can withstand

A 1/4 inch (0.025) aluminum tubing can hold 3,500 psi of pressure. The same aluminum tubing, but in 1/2 inch can only hold 1,800 psi. If we made a business jet pressurized fuselage out of the same tubing, it could only handle 182 psi.

When working with large pressure vessels, such as aircraft fuselage, don't be fooled by the low pressures. Because of their large size, these pressure vessels contain a lot of energy.

Don't forget the bulkhead. The bulkhead constains the fuselage skin, sucking up the load. A good lesson to learn is why the rear bulkhead failed on Japan Airline Flight 123.

Tubing and hose can be thought of as cylindrical thin-walled pressure-vessels. *

The strength of thin walled pressure vessels is determined by:

1. The material strength

2. The wall thickness, and

3. The size of the tubing.

The formula is: strength, psi = yield*(wall thickness/radius)

This last item, tubing size, is unusual. One can understand how strength is related to how strong the material is and how thick it is but size (radius)? The relationship between tubing size and strength is inverse; the larger the tube diameter the less strength it has. When you look at pressure ratings for tubing and hose you will notice that for the same hose, maximum recommended operating pressure goes down as the size goes up.

You can use this property to your advantage. For example, you might have a choice of tubing or hose size for a particular application. Everything else being equal, a smaller diameter line holds more pressure than a larger diameter line. Another advantage is that a smaller size weights less.

Inspection:

When you inspect a hose or line, you are inspecting a pressure vessel. As with all pressure vessels, they should be protected from damage that reduces the wall strength. Inspect for nicks, cuts, chafing, and corrosion. Make sure that the line does not vibrate.

*Pressurized aircraft are also pressure vessels.

Bulkhead Hole Size

2011-01-07T07:22:00.000-08:00

AN837 Bulkhead Fitting

Bulkhead Hole Size Chart for AN Fittings

Here is how bulkhead-hole-size is determined. Use the thread size as the outside diameter of the fitting. For example, a -4 bulkhead fitting has a 7/16-20 screw thread. The hole needs to be slightly bigger than 7/16 inch. A AN960 washer designed for a 7/16 thread has an internal hole size of .453 inch. So we make the bulkhead hole size the same size as the washer internal hole size.

MS21344 is used as a guide for installing bulkhead-fittings. The instructions to the left are from TO 00-25-223 INTEGRATED PRESSURE SYSTEMS AND COMPONENTS (PORTABLE AND INSTALLED). Slight differences between the two documents, specifically in the type of washer used. MS21344 indicates AN960 washer while TO 00-25-223 shows AN901 washer.

The discussion below concerns using bulkhead fittings in hydraulic systems. Note that this is not a good practice and generally prohibited by the military. "Universal fittings conforming with MS33515 and MS33657 shall not be used in boss applications in hydraulic systems..." MIL-H-5440G. Nevertheless, it is used and the bulletin below details problems with this practice.

Special-Airworthiness-Information-Bulletin-SAIB CE07-46

The FAA received reports of leaking hydraulic fluid due to improper installation of bulkhead universal fittings when installed in a hydraulic pump pressure port. The bulkhead universal fittings were turned in too far or not far enough causing the o-ring to contact the fitting threads resulting in o-ring damage and failure. While superseding standards and specifications exist, this SAIB refers to the installation of standard parts that resulted in the reports of leaking hydraulic fluid. The standard design installation of a bulkhead universal fitting into a port includes specific procedures to assure that the fitting is positioned so that the o-ring is located between, rather than on either of the two threaded portions of the universal fitting. These installation procedures are applicable, unless superseded by the Instructions for Continued Airworthiness for a specific airplane.

Standard designs for installation of a bulkhead-universal-fitting (flared, flareless, and straight threaded connectors) into a port utilize an AN6289 nut with a recess for a back up retainer for the o-ring. The use of an AN924 nut should no longer be proposed as a standard design for a new or modified installation of a bulkhead universal fitting into a port. The use of an AN924 nut instead of an AN6289 nut with a backup ring was initially included within the standard design per AND10064 (for flared tube and straight threaded connectors) for fuel and engine oil applications only. The use of an AN924 nut on a bulkhead universal fitting installed in a port became inactive for design in 1955 via AND10064. Refer to the attached excerpt from AND10064. The use of an AN924 nut instead of an AN6289 nut with a backup ring was initially included within the standard design per MS33566 (for flareless tube and straight threaded connectors) with nominal use identified for aircraft engine fluid connections. The use of an AN924 nut on a bulkhead universal fitting installed in a port became inactive for design in 1975 via MS33566. Refer to the attached excerpt from MS33566.

Previously approved installations using an AN924 nut that have acceptable in-service performance remain approved and remain acceptable. While acceptable performance of the AN924-nut on a bulkhead universal fitting installed in a port in low pressure hydraulic systems is known to have been achieved, un-acceptable performance in medium or high pressure systems is expected. Refer to ARP-4752 Aerospace – Design-and-Installation-of-Commercial-Transport-Aircraft-Hydraulic-Systems and AS4716 (R) Gland Design, O-Ring and Other Elastomeric Seals for additional seal information and general rule information that o-rings operating above 1500 psi should utilize backup rings. Installations of an AN924 nut on a bulkhead universal fitting installed in a port without acceptable in-service performance warrants review and consideration for a design change.

MS21344 installations and MS33566 installations (with AN6289 nut with MS28773 backup retainer) of a bulkhead universal fitting in a port is accepted by the FAA as a standard design for fluid pressures up to 3000 psi. Fitting design evolution continues. AS33566 retains the use of bulkhead universal fittings with an AN6289 nut and MS28773 retainer into a port and consequently is an accepted standard design. The FAA has also received reports that AS5440 includes information for the design authority to preempt the use of bulkhead universal fittings due to their problematic service history. Refer to AS5440.

Engine Balance and the Arms Race

2011-01-05T13:52:00.000-08:00

Interesting question about the engine balance "arms race" I've shortened and edited it a little:

So here's my question. Is there any computation or formula that you know of to convert the CH IPS velocity units into the moment units of fixed balancing machines, something that factors in the approximate weight of the whole engine and/or prop?
Here's what prompts the question. Some engine shops now balance engine crankshaft assemblies, and it's sometimes sort of an advertising arms race to split hairs more finely by advertising or claiming the lowest unbalance limit (expressed in moments). Sooner or later, some owner will ask, "OK, your crank balance limits are in different units from the units in the prop balancing book I read. How do your levels of engine balancing precision compare to what is acceptable or recommended for the prop?"

------------------------

This is not in my area and I don't know a thing about standard practices in the balancing industry, but I can't resist: Reducing the problem to its most basic level:

What Roger calls Arm is the eccentricity of the mass or the distance between the center of gravity and the center of rotation. So in terms of unbalance it is simply the amount of mass eccentricity times the radius (mR).

How do we detect this? We can spin the object and measure the vibration force because when mR >0 the centrifugal forces are greater than zero. We can measure ips, acceleration, g., or pilot comfort or whatever vibration measurement we wish to take. BUT these are all reactions to mR being greater than zero. We have measured the effects of imbalance and not the amount of imbalance itself.

Now Rogers question is astute. He wants to do just the opposite: Having measured the vibration in ips, g's, or human comfort level he is asking for a formula to convert any of these back to mR. I could be wrong but I don't think this is possible without knowing the mass. For example, If you tell me you are experiencing a 1 g force I cannot compute your mass (weight in this case). If Roger knows the mass of the crankshaft (and possibly the rpm, then I believe he could). He could empirically by making changes to the mass.

So if we talk about the amount of imbalance or mass eccentricity and not the vibration caused by the mass eccentricity we can express this in terms of eccentric mass and radius. So if Roger tells his customers that he balances crankshaft's down to "twenty milligrams per millimeter" (1 grain of rice 1 mm from the center of rotation) he will completely confuse the customer and possibly bullshit his way to leader of the arms race. The IPS guy has no idea what the amount of imbalance is - he just knows the amount of reaction there is to the imbalance.

To add further confusion, everyone is assuming that the crankshaft is perfectly rigid which it is not; that is why longer crankshafts have "counterweights" better described as "tuned pendulum absorbers". The crankshaft locally is not balanced as the cheeks are not opposite so we get local reactions to that imbalance. A 4 cylinder Lycoming or Continental engine has unbalanced reciprocating forces that are greater than any rotating imbalance. So ultimately it's a marketing question involving human nature and gullibility and thus so should the answer be framed.

1 gram is approximately the weight of 1 drop of oil so as the crankshaft rotates it is covered in oil so at the gram or sub-gram level the eccentric mass is always changing and thus crankshaft "balance" under operating conditions is dynamic and no amount of fixed mass will compensate. But for marketing reasons we could assume that some oil pools in recesses or is always present in oil galleys. Why not spin balance the crankshaft with the oil galleys filled with oil or the crankshaft wet to determine its mass eccentricity in real-life conditions and balance accordingly. You could advertise a 'wet balanced" crankshaft.!

Any counter-weighted crankshaft will have eccentric mass at the pins (bifilar mounted counterweights) as different diameter pins are used to tune the counterweight. Possibly one could compensate for this and call it a "bifilar tuned wet balance". Wait there's more we can do: any crankshaft collects a patina, sludge, and carbon deposits that certainly weigh in excess of our grain of rice. Why not advertise a "carbon compensated bifilar tuned wet balance."

Enough of this nonsense -

Inspecting Aircraft Control Cable

2011-01-02T07:15:00.000-08:00

Is this aircraft cable cable wear acceptable?

Aileron control cable failure on a Boeing-737-3TO on takeoff at Seattle, September 27, 1997 just six weeks after the cable was inspected for wear. Must have been a failed inspection -- yes? Not so quick, the inspection was performed "by the book." The inspection technique and process was at fault, not the mechanics.

The inspection consisted of checking for visible wear (external wire wear). However, the NTSB found that the internal wires were 90% worn! Most notably was the loss of aileron control on another Boeing 737-100, Flight-1659. The NTSB found that existing inspection methods could not detect the breakage of 98 of the 133 strands in the cable! Did you detect the broken strands in the picture above? Here is another picture with the tension removed.

Same cable with tension released

The NTSB investigation found that using professional FAA approved maintenance inspection at the most professional level will not detect dangerous control cable conditions. The broken-strands were not detected using the prescribed method of drawing a cloth rag over the cable. Only until tension was released from the cable were the broken strands detectable. Thus the need to release cable tension to better detect broken strands.

What about measuring the external diameter? The other Boeing standard at the time was to replace a cable when the the diameter of any single wire was reduced by 40%. This is called an "external wear" inspection. However, what the NTSB found in Flight-1659 was that cables wear internally as the individual wires slide past one another. This internal wear is greater on stainless steel cables than on galvanized cables because the galvanizing acts as a lubricant and stainless steel is noted for galling. Therefore, a maximum allowable reduction in cable diameter specification needs to be specified in the maintenance manual.

Notice also that stainless steel "the galling steel" wears faster than galvanized steel. Hmm, maybe stainless isn't so good after all.

In the 737-3TO incident illustrates the need for a cable diameter specification. the "NTSB found that several locations where the overall diameter of the cable had been reduced without damage to the exterior cable surface, which the NTSB metallurgist characterized as indicative of internal-cable-wear. In some locations, the cable diameter was reduced by as much as 0.03 inches (corresponding to approximately a 30% reduction in cable cross-sectional area for a nominal 3/16 inch diameter cable.)"

Same cable tension released and bent

And then there is the Twin-Otter crash killing 14 passengers in Tahiti in August of 2007 from frayed stainless steel control cables. The poor wear resistance of stainless steel rope has resulted in death and destruction. More frequent inspections are required for stainless steel flight control cables. For more information on this subject reference: Special-Airworthiness-Information-Bulletin:-SAIB CE-01-30, July 11, 2001.

There certainly has been enough time for the airframe manufacturers to update their maintenance inspection processes for flight control cables to reflect the lessons learned by the NTSB.

Aircraft Hose Temperature Limitation Warning

2011-01-01T07:27:00.000-08:00

Firesleeve DOES NOT increase the temperature rating of the hose.

Its sole purpose is to prevent flame penetration for a short period.

Aircraft Fuel Hose with Firesleeve Jacket Above

This fuel primer hose was completely burnt. When the pilot primed the engine he squirted fuel into the firesleeve.

This rubber hose is rated for 250 degrees F (121 C) and was located 4 inches from the exhaust stack. Firesleeve DOES NOT increase the temperature rating. There were plenty of burnt deterioration on the firesleeve to alert the pilot during preflight or the mechanic that there was a temperature problem.

Antiseze for Aircraft Mechanics

2010-12-31T08:51:00.000-08:00

Field Inspecting Bonded Joints - Problems and Expectations

2010-12-30T08:23:00.000-08:00

This is the condensed version of an article entitled "ASSESSING ADHESIVE BOND FAILURES: MIXED-MODE BOND FAILURES EXPLAINED"

Field inspections of adhesive bonded joints cannot detect degregation of bond strength. Yet at the end of a "successful" inspection, the aircraft mechanic is to declare the aircraft safe to fly. Is this an impossible task?

Adhesives depend on chemical bonds formed at the interface between the adhesive and adherend at the time the adhesive is cured. If chemical bonds are strong, failure occurs through the adhesive; bond strength is high. If the chemical bonds are weak or degraded, failure occurs through or near the interface; bond strength is low.

For metals, hydration of the surface oxides by water is the most common cause of failure. For example, aluminum forms an oxide almost instantaneously when the pure metal is exposed to the atmosphere after etching or abrasion during the production process. A bond which is susceptible to hydration at the interface has short term strength that may be sufficient to pass certification and quality assurance tests. However, as time in service progresses and the interface gradually deteriorates; bond strength degrades and eventually fails even without any loads.

Many current tests for process validation are based on static strength. For bonds which are susceptible to hydration, the chemical bonds at the interface are initially strong. It is not until the interface has begun to hydrate that there is a measurable loss of bond strength. Hence, short term strength or fatigue tests cannot prevent the in-service bond degradation and adhesion failures.

Many structures pass certification testing and quality assurance tests, including NDI, therefore one could infer that these are sound structures. Yet these structures may be susceptible to hydration of the interface and subsequent failure in service.

In later service, there is a potential for the adhesive bonds to dissociate so that the oxides can hydrate. This creates an interfacial failure of the adhesive bond. Moisture absorbed by the adhesive is sufficient to start hydration, and paints and sealants are not an adequate measure to prevent hydration because they simply slow down, not prevent, moisture absorption.

Out in the field, NDI can only tell whether or not the bond has a physical defect, it can NOT determine the strength of the bond. NDI can therefore not detect the onset of bond strength reduction. Aircraft mechanics cannot assure the strength of bonded joints and it is up to the regulatory authorities and the manufacturer to recognize the limitations inherent in the inspection of bonded joints.

Replacing Spot Welds with Rivets

2010-12-30T06:50:00.000-08:00

I am a Mechanic’s Toolbox subscriber. I looked through it, but could not find any reference for replacing spot welds with regular 426/470 rivets. The area of concern is in non-structural applications, such as access doors and fairings.

I have also looked all over the internet and have even talked with a few structures people and found nothing concrete. The closest is one fellow who verbally stated that it is generally acceptable to replace a fastener with one that is stronger, such as replacing an AN bolt with an NAS bolt.

Do you have any suggestions or information?

Your question has been already asked and somewhat answered on this engineering forum. No direct answer but some concerns. See the following link:

http://www.eng-tips.com/viewthread.cfm?qid=42081&page=62

In regards to higher strength fastener replacement - I see more hole failures than fastener failures ("smoking rivet") so concentrate on the condition of the hole. This seems also to be a theme of the link above.

Don't Forget to Inspect the Rocker Arm

2010-12-26T09:21:00.000-08:00

Pitted Continental Rocker Arm Face

When you purchase replacement cylinders the rocker arms are not included. Often they are placed aside until the new cylinders arrive and then installed. The rocker arm pictured above may wear out the valve guide very quickly. We have seen instances of 50-100 hours until the guide is worn sufficiently to cause valve leakage! If you then send out the cylinder for repair and then place the rocker arm back onto the cylinder you will repeat the problem. Here is what the rocker arm is telling you:

Notice the pitting on the top edge of the rocker face in this picture? The rocker arm face is tilted in reference to the top of the valve stem. The face should lie flat if everything is in alignment. By "everything" I mean the following:

Rocker shaft boss
Rocker arm bushing
Rocker face
Valve guide boss
Valve guide
Valve seat boss
Valve seat face

Quite a list isn't it. If any one of these are mis-aligned then the face doesn't sit flat onto the top of the valve stem. Here's the quick tip -- After installing the rocker arm carefully look at how the rocker face is sitting on the valve tip. If it is resting flat then you know that all the items on the list above are in alignment. This is a quick, easy, inexpensive, and informative inspection -- the kind I like!

Rocker arm should be flat against valve tip.

What if it isn't flat? Then you have a problem; The valve guides will wear prematurely and the problem can be any of the above. The first thing to check is to see if the rocker face has been re-ground, "refaced". Refacing is an awful thing to do as it is often done by hand, and it ruins the geometry of the face. Even if alignment is retained the cycloid curve is often flattened and this causes the pressure on the valve tip to move off-center causing the rocker arm to push the valve stem into the guide.

You can swap a suspect rocker arm with one that is resting flat from another cylinder to check if the problem is the rocker arm or the cylinder. If the rocker arm is bad, replace it.

Another tip is to inspect the rocker arm socket for wear. This must form an oil seal with the push rod ball. It is good practice to make sure each push rod goes back into the same rocker arm. The wear surfaces know one another and are compatible. If you mix them up then the surfaces are strangers and must "wear in" and form compatible surfaces.

Inspect rocker arm socket and push rod ball

Hole Quality for Aircraft Mechanics

2010-12-24T10:23:00.000-08:00

ROTEC radio noise

2010-12-09T07:12:00.000-08:00

John,

I found your website and the MF3-A looks like what may do the trick on my experimental aircraft that utilizes a handheld ICOM A-21 aircraft radio. The magneto on the Radial engine is putting out an RF signal that interfers with the radio. It goes away when I ground the mag and operate off the electronic ignition (which replaces the 2nd mag). ROTEC people have no specific recomendations but I think you have the right product to solve this problem. Am I correct in my assumption? Is this the correct suppressor? The ROTEC uses a non descript magneto and is not a Bendix or Slick . I would like to order one if you think it will help.

The MF3-A is designed for certain Bendix magnetos - it may not work or it might make the magneto not work correctly in other applications. How about the rest of your ignition system? Are the leads and plugs shielded?

The all important ignition system utilises two auto type spark plugs per cylinder independently fired by both a single self-energized magneto and Hall-effect 12 volt electronic ignition system, virtually eliminating total ignition failure when used in tandem. Timing is fixed at 22 degrees BTDC.

Shield the P lead wire. That is your first project.

Sounds like it is working as designed unless it is unique to your airplane. Possibly Rotec does not anticipate the use of radios in aircraft? I would push on them until I got to engineering to find out if this is a design feature or a defect.

There is a reason why modern aircraft engine use metal jacked shielding on their ignition wires and large shielded spark plugs as long-range radio communication under all conditions is vital to air safety. The aircraft industry moved away from the Rotec design and "automotive' spark plugs in the 1940's. Unless the laws of physics have changed since then, the Rotec design is inexpensive but unfortunately rather old and obsolete.

Glazed cylinder and high oil consumption

2010-11-21T08:31:00.000-08:00

Morning John,
I am finishing an annual on a 1949 Stinson with a Franklin Engine. I was checking your website for photos of a glazed cylinder. I suspect after 50 hours break-in that three of the cylinders did not seat.(excess oil burn, three bottom plugs very oily, oil blowing out of breather.
I plan to do a borescope after the differential press check and then a crank case pressure check.
If they did not break in do you suggest any chemical to pour into the cylinder to break the glaze and try again?

Well maybe not so good morning - about your cylinder...
The only solution is to remove the cylinders and re-establish a proper ring finish. And I emphasis proper ring finish. By proper I mean the ability of the hone shop to measure RMS finish and cross hatch pattern. Otherwise it is a random process with random results.

I don't have any pictures of "glazed" cylinders and I doubt they would look very much different from a normal cylinder. Usually, the term 'glazed" is used rather indiscriminately to describe any cylinder where the oil consumption is above normal after the break-in period. It is usually presumed that the reason for high oil consumption and poor compression is because the rings did not seat, but there can be other reasons such as improper rings or cylinder bore distortion. Failure of a cylinder to form compatible wear surfaces "break-in" is typically caused by an improper surface finish. High viscosity oils and poor temperature control are other reasons why rings do not seat.
There is nothing you can pour into the cylinder to dissolve carbon - but carbon is a symptom and not the problem.

The picture below shows what a hone pattern should look line.

Hone Cross-Hatch Finish Geometry

Introduction to Aircraft O-rings

2010-10-30T08:14:00.000-07:00

Aircraft Vacuum Pump Troubleshooting

2010-10-30T06:01:00.000-07:00

Inside view of a dry vacuum pump. The rotor and vanes are made from carbon. Notice that the top surface of the rotor is shiny or wet looking. This pump broke apart because oil got inside the pump. The oil mixes with carbon dust to form a sticky paste that will seize the rotor causing it to fracture.

Spark plug anti-seize

2010-10-27T11:03:00.000-07:00

John: From your past publications: "Champion recommends using 2602 spark plug anti-seize. Use sparingly. Some also use C5-a copper anti-seize although we prefer the Champion product which is a water based graphite."

Champion Spark Plug Anti-Seize and AutoLite Anti-Seize contain graphite. I suspect that Lycoming suggests not to use a graphite-bearing compound is because graphite can weaken aluminum.

Lycoming quotes the following in their "Lycoming Flyer" (http://www.lycoming.textron.com/support/tips-advice/key-reprints/pdfs/Key%20Maintenance.pdf) that "It is helpful to use anti-seize or plain engine oil for spark plug threads starting two full threads from the electrode, but DO NOT USE a graphite-based compound".

What is the proper anti-seize to use on spark plugs?

In aircraft the proper product to use is contained in the maintenance manual so there is no debate. In the case of Champion spark plugs it is 2602. This is also what every large maintenance shop I have ever been in uses so I don't understand why this would be questioned.

C5a is not only not a product listed in any of the manuals that I am familiar with but it is a bad idea for two reasons:
1. Spark plug antiseize does not contain any metallic particles that can enter the combustion chamber and cause preignition. (way back when they used to use a mica antiseize which is a mineral that prevents seizing and is also non-metallic so is preignition safe). Small copper particles are not something that one wants in an air cooled combustion chamber.
2. Spark plug torque specifications are based on using the proper lubricant (2612 for Champion) and any change in lubricant will change the friction and the torque/tension relationship. Proper torque for C5a is unknown.

The Lycoming flyer recommendation was most likely written by Joe Diblin ("engine joe") some 30 years ago so is somewhat dated. Although I agree with Joe. Graphite and water is an awful spark plug antiseize. I debated about bringing back the BG mica antiseize and selling it as it is a much better spark plug antiseize but the torque issue prevents me from doing so.)

Are Aeroquip fittings usable on Stratoflex hose, and vice versa, are Stratoflex fittings usable on Aeroquip hose?

2010-09-23T07:20:00.000-07:00

It depends. From a functional or a legal view?

It has been common and accepted practice to intermix Aeroquip 303 and Stratoflex 111 hose and fittings. The reason is that both are made to the same military specification (MIL-DTL-8794 for the hose) and are considered identical. I believe both Stratoflex and Aeroquip "discourage" the practice as neither one has control over the performance of the final product.

Aeroquip 601/701 and Stratoflex 156 (lightweight outside steel braid hose) the same argument could be said however the practice is not as common (ingrained) into the industry.

Teflon hose products: Not done.

Note that I fall back on "common and accepted practice" as the only justification. I cannot point to any authoritative document but just past practice in the industry. This is a thin leg to hand on as all it takes is one FAA inspector to wave his magic wand and disallow the practice. Often "accepted" practices become "unacceptable" as soon as a problem arises. Personally, I have no issues with intermixing hose and fittings on 303 or 111 style hose. I am more cautious with the other hose types.

Is it best to mate them to their own product line up?

Yes it is best to mate them to their own product line as that is how they are performance tested. For example, I know that when I build a Stratoflex 124-4 hose that its ultimate minimum burst strength is 12,000 psi. I know this by testing the product at random intervals not to exceed 500 hoses. If I deviate and build the product using some other method or use other fittings I do not know if it meets specification because there is no testing data that has been done.

There is another reason - As a mechanic one wants to shift as much potential liability "blame" as one can. Mixing manufacturer's gives each manufacturer an escape clause to deny responsibility if anything goes wrong. Do it per the manufacturer instructions and then if anything goes wrong it must be their fault.

Lycoming cylinder barrel wear signs

2010-08-11T14:18:00.000-07:00

Here's some not uncommon cylinder barrels.

First notice I've placed a white piece of paper to reflect the light onto the barrel. Now you aren't looking down a dark hole.

In the next picture the yellow arrow points to a dark patch. Dark patches are almost always areas of corrosion pitting. Next look at the red arrow. Notice two things; the rainbow colors and the shinny appearance. The rainbow colors is a heat tint that forms at approximately 500 degrees F. You should not see any hint tinting in a Lycoming or Continental cylinder barrel. The heat tint tells us the barrel at this spot got too hot. Notice that the heat is local. The heat is caused by blow-by of combustion gas past the piston rings.

Here is a closeup of the dark patch clearly showing corrosion pitting and the heat tint below.

Corrosion pitting is a difficult call. If it is concentrated in one spot it will trap oil that will oxidize and glaze over creating a tan colored patch on the barrel that will hinder the proper operation of the piston rings. If the pitting is more general then usually it does not adversely effect the operation of the piston rings.

When you look at a pit you might have the impression that "it doesn't look too bad" or that it is "real small" but you are only looking at the entrance to a hole. Pits penetrate deeply beneath the surface and they weaken the barrel. Proceed with caution and consult the manufacturer for rejection standards when it comes to any corrosion on a structural surface. Remember a cylinder is a pulsating pressure vessel. All the torque that turns that propeller is due to the piston pressing against the cylinder wall for leverage. On the other hand if you start rejecting cylinders with pits you will probably ground half our entire fleet of aircraft.

Breaking Studs

2010-08-11T08:04:00.000-07:00

"How often have you heard of a broken cylinder hold-down stud? A crankcase thru-bolt? All such failures are closely related and add up to be one of the most frequent types of structural failure experienced in piston aircraft engine's today?" (written in 1953 and still true today!) So let us consider the principal factors producing this type of failure and the important role maintenance plays in its prevention.

Cylinder hold-down studs, crankcase thru-bolts, and connecting rod bolts are a few examples of critically stressed parts subject to alternating loads. The manufacturer always gives them a specific nut torque range which must be maintained to prevent failure. The minimum torque value is necessary to avoid fatigue failure and the maximum to avoid exceeding the tensile strength of the stud. The vast majority of these failures experienced during engine operation result from fatigue. The nut is too loose, not too tight. When a nut is tightened too much often the nut yields or the stud pulls.

To give a better understanding of stud and bolt fatigue failure, we wish to review some typical applications via a simple illustration. Take a rubber band and note that it stretches in direct proportion to the amount of pull up to its useful (elastic) limit. Any change in stretch shows a corresponding change in load; a constant amount of stretch shows a constant load. Now, wrap this band tightly around two pencils. Assume that they are being held together with a force of two pounds. If you try to separate them with a force of one pound, what happens? Nothing that one can see - the pencils don't separate; the rubber doesn't stretch. And the rubber doesn't "feel" the pull. Why? The rubber band was pre-loaded to a greater force than you applied. The one pound pull only reduced the pressure between the pencils (from two pounds to one pound) and the rubber doesn't know the difference. It will not stretch further until the pull is greater than two pounds. If the band is made of metal instead of rubber and a load exceeding two pounds is applied intermittently, it will fail eventually from fatigue. To prevent fatigue failure, the preload must be equal to or greater than the alternating load imposed.

This principle applies to the cylinder hold-down stud and the other examples mentioned above. Consider the crankcase cylinder pad as one pencil, a portion of the cylinder base flange as the other, and the stud as the band. The flange gets a terrific tug at every combustion chamber explosion but the stud should not feel it - the design pre-load is greater than the pull of the flange. If the stud does feel it, the actual pre-load is lower than the design pre-load. Therein lies the story - "Another broken stud." The solution to this problem sounds simple. Just apply and maintain proper pre-load. But in actual practice a few complications arise. First, the torque method, which is far from fool-proof, is used to obtain a pre-load on the stud. But will all the cylinder hold-down studs get the same pull if each is tightened the same amount? Only if all conditions that resist the nut from turning are equal - thread lubrication, condition of the threads, condition of the mating surfaces, etc. It is important that they be kept as uniform as possible by giving close attention to the physical condition of the mating parts and abiding by the pertinent engine manufacturer's recommended procedures. The torque method has its drawbacks, but it is much better than guesswork and the best method generally available in the field.

The pre-load must remain unchanged. But, if the mating surfaces are not sufficiently hard (as in the picture above) and smooth, they can become indented during engine operation from the pressure exerted by the pull of the stretched stud or bolt. We have discussed how a stud or bolt stretches in direct proportion to the pre-load applied. The stretch also will vary in direct proportion to the effective length of the loaded part. For a cylinder base stud, which is relatively short, the actual amount of stretch is very little - somewhere in the neighborhood of .002 inch. What happens if its nut sinks .001 inch into the cylinder flange during engine operation? The pre-load is cut in half. The stud receives an alternating load many times a second. There is one chance of survival- a continued collapse of the mating surface more than another .001 inch. Then the nut will be finger-loose due to the slight clearance between the nut and the flange (except for intermittent contact permitted by bending of the flange), and the troublesome alternating load will vanish. Of course, other studs now will be overloaded and the cylinder base flange's load will be lopsided. The agony may be prolonged but eventually something is going to give. One of the biggest culprits affecting cylinder base stud breakage is the practice of using goop and sealants under the nut or flange.

This discussion has dealt with the man with the wrench. The integrity of an engine is very dependent on the man who puts it together.

This article is from "Engine Conditioning Summaries" produced at McClellan AFT in the 1950's. The problem of improperly torquing cylinder base studs has not changed in 60 years.

Lycoming Valve sticking Tip

2010-08-10T07:09:00.000-07:00

Valve stuck in open position - check for too little valve tappet clearance

If you have a sticking valve along with everything else you need to do to fix the damage and correct the problem, don't forget to check dry tappet clearance. Too little clearance can lead valve's sticking in the open position. Here is a link to how to check dry tappet clearance on a Lycoming engine.

Backfiring

2010-08-09T07:53:00.000-07:00

Under certain conditions of mixture ratio there will be backfiring in the:

intake manifold,
exhaust manifold.

Backfiring in the intake manifold or carburetor: Most frequently occurs during starting of an engine under cold-weather conditions. The priming and choking operation varies the mixture from too lean to too rich. A very lean mixture will burn very slowly and the charge may still be burning when the exhaust valve is closing and the intake valve is about to open. The fresh charge in the intake manifold is not so diluted as when induced in the cylinder and mixed with the clearance gases and consequently will burn more rapidly than the charge in the cylinder. If the fresh charge, upon being induced, is ignited by the residual flame of the previous charge, the flame will travel back through the intake manifold, burning the charge therein.

Backfiring caused by the slow flame propagation of a lean mixture is not confined to starting, but may occur under any condition of engine operation if the mixture becomes lean enough; it can be made to occur by excessive leaning of the mixture with the mixture control. Backfiring in the carburetor during starting conditions only occurs when the mixture is too lean. Rich mixtures burn faster than lean ones, and under starting conditions the extra fuel which must be supplied to form a rich mixture is probably partially evaporated by the heat of combustion and extinguishes the flame before the next charge is induced.

Backfiring in the exhaust system: Backfiring occurs in the exhaust system under two conditions of operation. The most common occurrence is the somewhat irregular backfiring that occurs when the engine is being motored (driven by the propeller) with the throttle closed. Sometimes you can hear this backfire on airplanes on short final to land. Under this condition the idling system is supplying the mixture. The manifold pressure will not vary much with speed so that the quantity and quality of the mixture in the manifold are practically the same as under normal idling speed. Also, the exhaust product in the clearance space remain the same with speed., consequently as the engine speed is increased due to motoring the amount of the charge per stroke becomes smaller and the dilution greater until firing ceases. The succeeding unburned charges are pushed out into the exhaust system, the dilution in the clearance space is decreased and after a few cycles with no firing, a charge will be fired. It will be a lean and slow burning charge and the opening of the exhaust valve and result in an explosion. This type of backfiring can be eliminated by increasing the richness of the idling mixture.

The other condition that results in backfiring in the exhaust system is usually that of a faulty fuel control. Under part throttle operation, a faulty carburetor may cause an enrichment of the mixture which would cause misfiring. Opening the throttle would reduce the richness, and the firing of the charge would be resumed. In the meantime, the unburned mixtures which have collected in the exhaust system have become combustible probably due to the condensing of some of the heavier ends of gasoline, and these are ignited from the flame of a cylinder which fires resulting in a rather violent explosion.

Faulty NTSB Conclusions N9348S

2010-07-16T16:56:00.000-07:00

I have to stand up for the mechanic as this NTSB report's "probable cause" is just plain stupid.

NTSB Identification: CEN09LA209
Accident occurred Sunday, March 15, 2009 in Bellefontaine, OH
Aircraft: BEECH B24R, registration: N9348S
Injuries: 2 Uninjured.During cruise flight the pilot noticed abnormal engine noises and a partial loss of engine oil pressure. He immediately diverted to the nearest airport, but during the turn to base leg the engine oil pressure dropped to zero pounds per square inch and the engine seized. The airplane was not in a position to reach the runway threshold or to clear the airport perimeter fence. During the landing rollout the airplane impacted the airport perimeter fence, damaging both wings and the nose landing gear. An engine teardown examination revealed that the Number 3 cylinder connecting rod assembly had separated from its corresponding crankshaft journal. The journal surface was blue in color, consistent with exposure to excessive heat and lack of lubricant. The oil suction screen was obstructed with bearing material. The Number 3 cylinder connecting rod cap was found jammed beneath the counterbalance weight. One of the two connecting rod stretch bolts remained intact. The corresponding nut was found finger tight. The measured torque for the Number 2 cylinder connecting rod bolts were significantly less than the manufacturer's specification. The engine had accumulated a total of 3,799 hours since new and 492.7 hours since its last overhaul in 1999. The engine was last inspected 23.9 hours before the accident occurred.
The National Transportation Safety Board determines the probable cause(s) of this accident as follows:
The inadequate torque of the Number 3 cylinder rod bolts by maintenance personnel, which resulted in a failure of the connecting rod and a total loss of engine power.
---------------------------------------------

" The fallacy here is the idea that measured bolt torque today is the same as what was originally applied some time in the past? The ASTM Specification for Structural Joints sums up the problem: "The uncertainty of the effect of passage of time and exposure in the installed condition." I would add especially after the engine broke into pieces possibly stretching and bending the bolts that remained intact!

When using a torque wrench to break-loose a nut the torque wrench is measuring the amount of torque required to overcome friction.
1. A little bit of off-torque applied to the wrench might be just enough overcome friction and get the nut to turn if the twisting force was locked into the bolt.

2. Friction might have changed due to time or the forces of engine failure on the bolt. The NTSB took a friction measurement that has an unknown relationship to torque applied some time in the past.

3. The two "intact" bolts most certainly suffered from some unusual loading as the engine destructed. If they were yielded and stretched then of course the nuts would be loose.

4. Fretting of the faying surfaces is a classic example of how torque is lost due to surface wear. So the answer is no for 4 reasons.

Torque is a twisting force that tightens the joint not only by stretching the bolt along its axis, but also by twisting it. We stretch AND twist the bolt into a coiled spring. The stretch is necessary as it keeps our surfaces clamped together, but what is the twisting force doing? It might have gone into loosening our nut slightly the moment we released tension on our wrench; or it might still be locked into the joint in which case all of the wrench force you used to turn the nut may still be there trying to loosen the nut. It all depends on the friction between the nut face and the faying surface (an important lubrication consideration).

What happens if we use a torque wrench to “check the torque” on a bolt? If we torque the nut until we reach the “breakaway torque” we have to apply enough torque to overcome the friction between the nut face and joint surface. Until this happens none of our torque is felt by the bolt. Once the nut face is released our applied torque starts to twist the bolt until we apply even more torque until it overcomes the friction between the male and female threads. Only now does the nut move relative to the bolt and we detect the “breakaway” torque. Notice that the only thing we are measuring is friction. Everything depends on friction. If the friction is more or less than when the bolt was originally tightened then our “breakaway” torque will be more or less than the torque that the assembler applied to the nut. That is the quandary. Did the assembler improperly torque the nut or did the friction change? We have no way of knowing.

“The nuts were loose.” Can we check assembly torque by loosening the nut? Same problem but only worse because if the twisting force is locked into our bolt then it might take only the smallest amount of torque to get the nut moving (our breakaway torque) as the bolt tries to unwind.

But what if the nut really is loose? Then take your pick; all or some of the reasons listed below:

1.Embedment relaxation of the faying surfaces
2. Bolt stretch from metal creep (especially at high temperature)
3. Nut backed-off due to vibration loosening
4. Wasn’t tightened properly to begin with
5. Elastic interactions between multiple bolts in a flange has reduced preload (crosstalk).

Checking the degree of tightness in a joint by using a torque wrench to measure the breakaway torque is not accurate and leads to incorrect conclusions. Beware of inspectors carrying torque wrenches.

" When bolt pretension is arbitrated using torque wrenches after pretensioning, such arbitration is subject to all of the uncertainties of torque-controlled calibrated wrench installation that are discussed in the Commentary to Section 8.2.2. Additionally, the reliability of after-the-fact torque wrench arbitration is reduced by the absence of many of the controls that are necessary to minimize the variability of the torque-to-pretension relationship, such as:
(1) The use of hardened washers;
(2) Careful attention to lubrication; and,
(3) The uncertainty of the effect of passage of time and exposure in the installed
condition." quote from Specification-for-Structural-Joints Using ASTM A325 or A490 Bolts

Lycoming camshaft

2010-07-16T14:44:00.000-07:00

Got a IO 360 Lycoming L2A in a Cessna 172. Approx 500 hours since last overhaul by a major facility. Cam was not replaced at last o/h and at that time engine had run 2480 hours. They fitted tappet bodies p/no 72877r ohc which I take to be overhauled. We have ferrous metal in main filter, enough to cover your little finger nail. Do you think we are over reacting going for warranty? I haven’t pulled a jug yet. There was also a small piece of metal which looks like the tail of a cotter/split pin.

--------

No I do not think you are over-reacting. Something inside the engine stinks and it ain't getting any better.

The principle stress on cam and cam followers is below the surface. Fatigue failure is cumulative and starts below the surface as a crack which eventually reaches the surface and releases a flake. Inspecting and resurfacing "reconditioning" the surface does not restore the fatigue strength. Failure is only a matter of time - total time. Zero time requires that you go back and melt the steel and start over.

Camshaft visual inspection guide

Setting Slick Magneto Point Gap

2010-07-13T08:06:00.000-07:00

.008 to .010 inch for Slick 4300/6300 series magnetos. But Wait! You do not set point gap. You set E-gap. On the Slick magneto, E-gap is critical; not point gap - much easier starting! One establishes E-gap (point opening at defined rotor position) and then point gap should be within the limit range.

Shown here is the E-150 E Gap Tool being used to set rotor position. Then use your timing light to detect point opening. By point gap we mean the maximum distance the contact points separate. Reference the Slick Maintenance manual for complete instructions on setting internal magneto timing.

Slick E gap timing tool(red arrow) inserted into slot in rotor and against poll shoes. Notice how magneto is oriented with the coil up. Locate the appropriate L or R timing slot (shown in lower picture) on the rotor magnet and insert the notched end of the T-150 into the L slot for Left-hand rotation magnetos and the R slot for Right-hand rotation magnetos.

Place tool on Left side for L magnetos and Right side for R magnetos. Rotate the magneto clockwise for Left-hand rotation and counterclockwise for Right-hand rotation magnetos until the T-150 rests against the pole laminations. The magneto is now at the E gap position. Adjust the points to be just opening and you have set inernal magneto timing. Point gap opening should be .008 to .012 inches. Torque adusting screw to18-20 in-lbs. Torque the pivot screw to 15-18 in-lbs. Note: many mechanics do not use this simple tool. They estimate E gap. Comparing this method with the tool shown above you can be off considerably. This may account for differences in rpm drop between magnetos when doing the magneto-check.

T-150 Slick E-Gap Timing Tool

Thickness: .060 inch
Length: 2.90

Lycoming Fluctuating Oil Pressure

2010-06-18T15:27:00.000-07:00

Below is a description of the problem that is quite lengthy. Sum it up to say that the oil pressure keeps changing. What could it be?

Here is my take on all of this: If you re-label your oil pressure gauge to say "valve opening and closing gauge" then does this better describe what you are seeing on the gauge? Knowing nothing about your engine but having a good seat for the ball seems obvious.

Engine: IO-360A3B6D
Installation: 1980 Mooney 201 (M20J)

Time since overhaul approximately 100 hrs.
Time since overhaul approximately 24 months.
Always in a hangar.

Approximately 6 months ago I started to notice a gradual decrease of oil pressure once in flight. Start-up pressure was good, take-off pressure was good, initial cruise was good, but shortly thereafter it would start to drop.

Where I would see pressure at approximately 72 psi during flight, it gradually became 70, them some time later during a subsequent flight it would 68. That eventually became 66, then 64, then low 60’s and even 58.

By the time it hit the low 60’s I began to adjust the oil pressure relief valve and each time I adjusted it, the initial (start-up) pressure re-established itself at around 90 to 94 with cold oil and in flight I would see 3 or 4 psi more than the previous flight. Generally I would be able to re-establish mid to high 60’s for a flight or two before it would work it way down to the low 60’s. Eventually I ran out of adjustment of the oil pressure relief valve adjustment screw (spring + ball).

One weird characteristic to mention is that the oil pressure would increase on approach to landing despite the lower power setting on let down to landing. If I had 60 psi at the point of flight just before let down with perhaps 23” HG and 2450 RPM, I would see 64 psi at 15” HG and the resulting RPM.

With the spring adjustment now bottoming out, I changed the spring and ball. I installed the next heavier spring then went flying and experienced exactly the same problem. One thing we noticed was that the oil pressure relief seat didn’t appear to be round thus not allowing the ball to seat properly. I will be getting the tool to have this corrected. If I read your oil pressure problem check list it mentions that if this problem is experienced then it is likely that the result would be low pressure at low rpm. It is possible that this could affect the pressure at cruise as well or just by virtue of the way the ball naturally moves away from the seat at higher rpm means that this can’t be the problem?

Other items that have been mentioned are the oil cooler (but it was good at the time of the overhaul of the engine, the gauge (I have 2, and a digital E.I. and the ships gauge) and both seem to give the same readings. Both are hooked up to the same oil line. Someone else mentioned the veratherm (spelling?), the oil pressure filter by-pass valve, the oil quantity which I usually keep at 6 because I fing that at 7 it blows it back to 6.

By the way, unless I have a bad oil temp gauge the oil temps are normal (mid-green).

If you have any ideas it would greatly appreciated.

Should Corrosion Preventive Fluids be used on Riveted Joints?

2010-06-08T07:38:00.000-07:00

Spraying a corrosion inhibitor into the wings of aircraft results in it's leaking out at seams and rivets and sometimes creating a 'smoking" rivet" appearance. Two possible explanations for the "smoking rivet" are:

1. the lubricant has carried dirt and debris to the outside surface or,

2. the wet skin around the rivet and seam attracts dirt.

Some mechanics have suggested a third possibility: that in highly loaded riveted lap joints (such as the Cessna 310 wing outboard of the nacelles), the lubricant might be creating a loose joint. In other words the lubricant is creating the "smoking rivet by interfering with the load transfer of the joint. I have done a literature review to see if this idea has any merit.

Transfer of the applied load through the joint is shared between frictional action at the faying surface and contact between the hole boundary and rivet . Lubricating the faying surfaces might create partial slip and transfer all of the applied load to the rivet. Slip also creates fretting and this leads to cracking. (Farris, T. N., Szolwinski, M. P., and Harish, G., “Fretting in Aerospace Structures and Materials,” Fretting Fatigue: Current Technologies and Practices, ASTM STP 1367).

"Comparative flight-simulation tests on four types of joints of 2024-T3 sheet material were carried out with and without application of the penetrant LPS-3. Types of joints were double strap joint with hi-lok bolts, asymmetric strap joint, lap joint with countersunk rivets and lap joint with dimpled holes. Observations on slip during the fatigue tests were made, also in static tests of failure. A reduction of fatigue life was observed in two joints. The effect depends on the design of the joint, while the maximum load in the test may also be significant in view of the occurrence of slip." Effect of an Anti-Corrosion Penetrant on the Fatigue Life in Flight-Simulation Tests on Various Riveted Joints Schijve, J | Jacobs, F A | Tromp, P J Natl. Lucht. Ruimtevaartlab. Vol. NLR TR 77103 U, pp. 34. 31 Aug. 1977

"Some aircraft manufacturers and operators have attempted to control in-service corrosion by the use of water-displacing organic inhibitors which can be either brushed or sprayed onto corrosion-susceptible areas of the structure. However, because of the low surface tension and lubricating properties of these preparations, concern has been expressed as to their potential side-effects on the fatigue performance of bolted and riveted joints. Fatigue tests were carried out in repeated tension under both constant-amplitude and multi-load-level sequences on several types of 8-bolt double-lap joint specimens of 2024-T3 alclad aluminium alloys sheet.
Tests were made on joints assembled with either ‘dry’ components or components coated with the corrosion inhibitors LPS-3 or PX-112. Contrary to the findings of previous investigations into the effect of inhibitors on riveted joints, the two corrosion inhibitors used were found, in general, to have either no effect or a beneficial effect on the fatigue lives of bolted joints. It is concluded that the specific effects of a water-displacing organic corrosion inhibitor on fatigue strength of joints are likely to be dependent on the type of joint, its configuration and on the severity of the load spectrum involved." Water-displacing organic corrosion inhibitors—their effect on the fatigue characteristics of aluminium alloy bolted joints A.S Machina and J.Y Mann

"The results showed that the lives of the treated shorter than those of the untreated specimens." A Short Study of the Effect of a Penetrant Oil on the Fatigue Life of a Riveted Joint by P. l-f. O’Nei/l and R. I. Smith Structures Dept., R.A.E., Farnborough

"It is concluded that joints fabricated with lubricated rivets, like those fabricated with lubricated threaded fasteners, have lower bearing yield and bearing ultimate joint strengths than when fabricated with clean fasteners." EFFECT O LUBRICATION ON THE JOINT BEARING STRENGTH OF RIVETED LAP JOINTS DEPARTMENT OF THE NAVY NAVAL AIR DEVELOPMENT CENTER AIR VEHICLE TECHNOLOGY DEPARTMENT REPORT NO. NADC-72055-VT

" It was found that in joints with high fastener-clamping force, the application of lubricative corrosion-prevention compounds increases the fatigue life, whereas the use of CPC is detrimental to the life of joints with low fastener-clamping force." Fatigue behaviour of aluminium alloy 7075 bolted joints treated with oily film corrosion compounds

"The results showed that the presence of CICs had a significant influence on the fatigue life, and also on the failure mode of the joints. At high load levels, the application of CICs caused a reduction in the fatigue life of the joint by more than a factor of two. In this load range, the CICs appeared to cause the failure mode to change from tensile failure of the sheet (the prevalent mode at medium load levels) to shear failure of the rivets (observed at the highest load levels). Specimens that failed by rivet shearing showed some fatigue cracks propagating along the critical rivet row. In treated specimens tested at medium load levels, a reduction in the fatigue life still occurred, with all specimens failing in the sheet. At low load levels, there was little difference in fatigue life for the three conditions, although specimen test run-outs meant that further testing will be needed.

The results are believed to have significance for managing the small aircraft in which these joints are common." Corrosion treatments and the fatigue of aerospace structural joints Aditya Jaya Ung Hing Tion , Reza Mohammed, Cees Bil and Graham Clark School of Aerospace, Mechanical and Manufacturing Engineering, RMIT University, Bundoora, VIC, 3083, Australia

Factory New Limits and other Nonsense

2010-05-30T10:19:00.000-07:00

Pure nonsense! I can give you an example of how one engine manufacturer's cylinder barrel new limit is the same as maximum service limit. What... new limit=service limit!

But there is a story behind this so here goes...

Many years ago one of the engine manufacturer's had a problem with excessive cylinder barrel wear on a high-performance airplane model. There was a meeting of the user group for that airplane and there the factory rep assured all of the owners that they would take care of the problem under warranty. They would boroscope the cylinders and any cylinders that had excessive barrel wear (no cross-hatch left i.e. bore polished) would be removed and re-honed.

Fair enough, but some of the owners were concerned that they were going to get back cylinders that were not to new limits and close to being worn out. The factory rep assured them that "any cylinder not in new limits would be replaced with new". More than fair.

So everyone's happy and the factory rep goes to my shop and says "John, we want you to do all of the cylinder work on this warranty problem." OK, I say, "but if the barrel is already worn out and I hone it, it will be larger. I can't hone it smaller so I know they won't be within factory new limits."

John, the factory rep says, you don't understand. You are going to hone it to factory new at +5 oversize. But Mr. factory rep, you don't have a +5 oversize. We do now!

So here is how this works:
Mininum new is 5.000 inches
Maximum service limit is 5.005

Mininum new for 5 over is 5.005
Maximum service limit for 5 over is 5.010

So you see these cylinders were factory new but at +5 oversize, or were they service limit for standard size? Take your pick as they are both!

Now it's not all that bad because maximum service limit is not a wear limit as such. It doesn't mean that the cylinder barrel will stop working at that size. It is a repair limit or the limit of size during repair in which the engine will perform satisfactory during its anticipated life. All things being equal we want it be as close to new size as possible but there is way too much emphasis by users on "new limits".

Now that we're talking "new limits" or 'factory new limits" there is one more gotcha. We got into trouble above by not having a thorough understanding of new limit. Now we're in trouble because of imprecise thinking; marketeers in this industry take advantage of it (you).

If you specify "Lycoming new limt" or "Continental new limit" we can all agree on what this limit is as it is published in their "Table of Limits". But when we say "factory new limits" then who's limits are these? You might incorrectly assume that it is Lycoming or Continental. There are shops that work to their own limits and to them "factory" referes to themselves.

Increasing Reliability of your Aircraft's Fluid Delivery System

2010-05-30T07:48:00.000-07:00

Each hose assembly adds 6 fluid connections:

Between hose and fitting x 2
Between hose fitting and attachment fitting x 2
Between attachment fitting and component x 2

Each fluid connection becomes a possible point of failure.

Design your to system minimize fluid connections.
Avoid the use of adapter fittings where possible.
Consider trade-off's before purchasing popular upgrades such as fuel flow transmitters, remote oil coolers and filters, external engine breathers, etc.

Preventing Stainless Steel Screw Seizing, Galling, and Stripping

2010-05-29T09:37:00.000-07:00

Compass degauss

2010-05-26T09:29:00.000-07:00

John,

Just read about your airframe degauser, really interested!

I have a CubCrafters light sport cub, so much magnetism that the compass is always pointed N/E. Have you had any dealings with this model aircraft and if so what was the success?

Have tried to degause with an old modified TV degauser, helped some but not nearly enough. When checking for hot spots with a hand held compass we find a strong field in the firewall/engine mount area, haven't pulled the mags yet to work on this area.

Bill,

I have no experience with CubCrafters.

1. A compass reacts to any piece of iron so is not suitable for detecting artificial "non-earth" magnetism. The reason for this is that iron is always magnetized with the earth's field. Earth's magnetic lines of force find it easier to travel in iron than in air so the iron sucks in the fields and then sprays them out at any radius in different directions. The only method of solving this problem is to add distance between iron and your compass and use the compensator to correct for any remaining disturbance.

It is important to distinguish the two as you cannot degauss the earth's field. The best method without a suitable meter is to clip two paperclips together so one dangles from the other and come up close to suspected magnetized iron parts and see if the paperclip is attracted to the iron. (if it is then you must use a new paperclip as now your paperclip is magnetized and will stick to anything iron).

The key is to:
1. Use a meter to detect artificial fields, and
2. Use a degausser at that spot that is stronger than the artificial field, and
3. Decay the field in the proper manner.
4. Go back and check your work with the meter.

This has proven to be the most successful method over the past 15 years since I developed the system. In cases where it doesn't work I refund the rental fee.

Aircraft Structural Screws

2010-05-26T06:57:00.000-07:00

aircraft battery explosion

2010-05-07T18:10:00.000-07:00

Aircraft Battery Explosion - How to prevent

Hydrogen gas explosion. Note the battery is burned from the top down. This is a lead acid battery with vent caps - the kind you add water too. During charging explosive hydrogen gas is vented from the caps into the battery box. The box must be vented and have sufficient air circulation to prevent the build-up of hydrogen.

Also, rapid charging a discharged or "dead" battery generates large amounts of hydrogen gas. "Jump starting" a dead battery and then allowing the aircraft charging system to charge the "dead battery" can dump the full capacity of the aircraft alternator into the battery causing large amounts of hydrogen gas. One does not "jump start" an aircraft battery and then go fly the airplane for many safety reasons.

This battery blew the lid off when hydrogen gas was not properly vented. Notice several black caps on top of the battery have blown their tops off exposing the lead vent plug. To summarize:

1. Non-sealed battery's must be vented and have air circulation.

2. Never use the aircraft charging system to charge a dead aircraft battery.

Note: sealed recumbent gas aircraft battery's vent much much less hydrogen gas as most of it is recombined inside the battery. Both Concord Battery and Gill manufacturer Recumbent gas batteries.