In October 2014, the United States Air Force required a replacement to their legacy Honeywell and Sperry ADT-222B test sets. They had 20 test sets, 12 of which were deemed unserviceable and being used to cannibalize for working units. These units were integrated into the for air data controller for the Electronic System Test Stations (ESTS) system. The system is used to test aircraft air data computers and provide accurate pneumatic pressures to the aircraft air data computer system over a remote bus. The Honeywell assets had been manufactured in the 1980’s and had become nearly impossible to support due to obsolescence and available parts.

A competitive, full and open solicitation was released for a newer and supportable replacement test set. The primary objective of this program was to replace the ADT-222B, while providing a “plug and play” drop in replacement. This meant that the replacement had to be capable of communicating remotely over the IEEE-488 bus using the legacy Honeywell 1975 protocol.
Raptor Scientific (TestVonics, Inc.) proposed a derivative to our USAF approved ADC-2500 Series Air Data Calibrator, a low technical risk approach. The ADC-2522 was uniquely designed to support the ESTS ADTC program. Over the course of 14-month engineering evaluation, our team successfully developed and qualified the ADC-2522 to fully replace and emulate the Honeywell ADT-222B.

The ADC-2522 is a high precision air data test set calibrator and can precisely control and measure Altitude and Airspeed pressures. It meets and/or exceeds all the performance characteristics set forth by Tinker AFB for the Honeywell ADT-222B and ADT-222C. The ADC-2522 can be operated using local controls including the 10.4-inch touchscreen display with intuitive graphical user interface. It can also be operated remotely through RS-232 or IEEE-488. Because of this, the ADC-2522 is qualified to the same National Stock Number (NSN: 4920-01-105-0016) as the Honeywell ADT-222B (P/N: 4047505-421).

Since 2016, twenty (20) of these units have been in continuous operation in USAF facilities. There have been zero reliability issues reported and the USAF has full capability to support and calibrate these units using existing calibration standards.

Since then, our ADC-2522C and ADC-2550V3 derivatives have been used to replace the Honeywell ADT-222B and ADT-222C. They have the same ranges and accuracies and feature additional interfaces like RS-232, USB, VGA and RJ-45. The most notable difference between the ADC-2522 and the ADC-2550V3, is that it does not include the Honeywell Sperry IEEEE-488-1975 interface and software emulator. Customers replacing their test set with the ADC-2550V3 are opting to update their ATE to command remotely over IEEE-488.2 or RS-232 using our SCPI communications protocol.

• ADC-2550V3 Air Data Calibrator

Looking for a replacement for your Sperry or Honeywell ADT-222A, ADT-222B or ADT-222C test set? Contact Raptor Scientific for more information.

If you would like to download a pdf version of this case study, please submit the form below.

The post Case Study: United States Air Force Honeywell ADT-222 Test Set Replacement appeared first on Raptor Scientific.

Below we will cover some answers to questions that are commonly asked by avionics technicians.  This article is intended to help technicians testing aircraft pitot-static systems the basic air data principles and how to use the ADTS to troubleshoot them. Some air data principles can be confusing.

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If I only changed an aircraft altimeter, why would I have to connect the pitot hose as well as the static hose to perform a check?

If there are any differential pressure-sensing instruments (Mach indicator, airspeed indicator, air data computer, etc.) connected to the same static source as your altimeter, you MUST connect the ADTS pitot hose to the aircraft before running up the static channel.

Remember that the airspeed indication is a function of differential pressure or Qc. Qc is the difference between Pitot pressure (Pt) and Static pressure (Ps). Qc is calculated using the following simple formula: Qc = Pt = Ps

Zero knots of pitot pressure is always equal to the static pressure at any altitude.  For example, let’s say your ambient pressure is 30.000 inHg.  This means that the pitot pressure for 0 knots is also 30.000 inHg.  If you control the ADTS up to 10,000 feet, the Ps is reduced to 20.580 inHg, and the pitot pressure required to make your airspeed indicator read 0 is also reduced to 20.580 inHg.  If the ADTS is not connected to the pitot system, it will obviously not be able to control the pressure in the pitot system.  The 30.000 inHg of outside pressure will be in the pitot system causing the airspeed indicator to increase as the ADTS altitude increases.  Remember, at 10,000 ft., 200 knots is approximately 22.500 inHg.  If 30.000 inHg is in the pitot line at 10,000 ft., there will be a Qc of nearly 10 inHg which will cause the airspeed indicator to increase to over 400 knots.  At 30,000 ft. (8.880 inHg) the Qc will be approximately 23.000 inHg which will cause the airspeed to read over 600 knots. If your airspeed or Mach indicator is not capable of going to these speeds, you will most likely damage them.

You would never intentionally apply 600 knots of pressure to an airspeed indicator that only goes to 400.  In order to avoid doing it by accident, it is important that you understand how airspeed indicators and the ADTS operate.  Always use both pitot and static hoses when you perform maintenance with the ADTS.

While using the ADTS, I saw an airspeed indicator still reading correctly with an “out of tolerance” pitot leak.  Why is this?  Are pitot leaks really that important?

Don’t confuse how the ADTS works with the way an airspeed indicator or an altimeter will function in flight.

The ADTS automatically compensates for small leaks by using its pumps to maintain whatever airspeed or altitude you command.  When you perform a leak test, the pressure sensor within the test set and the unit under test are both isolated from the output of the pumps and the ADTS will then only monitor the pressure in the lines.  If there is a leak, you will see the pitot or static pressure try to equalize with the pressure outside the lines

In flight, there are no pumps to maintain the airspeed or altitude at the proper indication.  Any leak will cause a pressure change in the pitot system and therefore an erroneous airspeed reading.

During takeoff (when the cabin is at ambient pressure), a pitot leak will cause a low airspeed reading.  As the aircraft climbs (and the cabin is pressurized), the airspeed reading will increase due to the higher cabin pressure entering the pitot line.

Altimeters will always read low if the static system leaks inside the pressurized portion of an aircraft.  The greater the leak, the more erroneous the reading.  A leak-free pitot-static system is very important to the safety of flight.

I noticed that when I do a pitot leak test, the airspeed reading will sometimes increase rather than decrease.  Why is this?

This is probably because you did the leak test while the altitude was controlling up. Typically, when both pitot and static leak tests are required, technicians may control both the pitot and static simultaneously in order to save time.  This is where the confusion starts.

Try not to think of a leak test in terms of feet or knots—think of it in inches of mercury (inHg).  Let’s say the outside pressure is 30.000 inHg.  Your technical order might instruct you to perform a pitot leak check at 200 knots.  This means that if your ADTS altitude is at field pressure elevation, the pitot pressure will be approximately 31.900 inHg (greater than the outside pressure).  If you perform your leak test at these settings, a leak will appear as a decrease in airspeed.

Now let’s say that your ADTS altitude is at 10,000 feet (20.580 inHg).  This means that the same 200 knots of pitot pressure is now only about 22.500 inHg (less than the outside pressure).  If you do your leak test with the ADTS at these settings, a pitot leak will appear as an increase in airspeed rather than a decrease because the 30 inHg on the outside of the system is leaking into the pitot system causing an increase in pressure.

This is important to know because there is an altitude where 200 knots will equal your outside air pressure.  This altitude is only about 2,000 to 4,000 feet above a

mbient.  (refer to Figure 2.)  If you do a leak test at that altitude, you will have a no-leak indication even if your hose is wide open because there is no differential pressure between the pitot system and the outside air.

You should always do your pitot leak tests while the ADTS altitude is at or below field pressure elevation in order to maximize the pressure differential

How do I know what altitude to set in the ADTS?

It’s always good practice to know the pressure altitude (ground) prior to connecting the ADTS to the unit under test, put the unit into measure mode and open the front panel ports to ambient. You can open the manual vent valve as well.  Additionally, the ADTS features a course barometric sensor that “captures” the ground altitude when the system is powered on.

Always enter in the pressure elevation.  Remember that field pressure elevation is NOT actual elevation.  This is a very common mistake that you will see a lot. If your actual elevation is 200 feet above sea level, your pressure elevation will most likely be different depending on local weather conditions.  On a cold day with a very high barometric pressure, your pressure altitude may be less than -500 feet below sea level.  On a hot day with a very low barometric pressure, your pressure altitude may be 900 feet above sea level or more.

What does it mean when you get a track loss condition?

The ADTS is capable of compensating for small leaks through the use of vacuum and pressure pumps.  When the pressure leaks out of the system at a rate faster than the ADTS is able to compensate, (when there is a wide open pitot or static leak, for example), the track loss light will illuminate and the ADTS will go into the reset mode.

When this happens, open the vent slowly and allow the pressure to stabilize before reapplying pressure.

When should I perform an auto exercise on the ADTS?

The general rule of thumb was to perform the auto exercise whenever the test set has been turned off for more than 4  or more hours, or when power has been on for more than eight hours without changing airspeed or altitude settings. Auto exercise helps to run the transducers fully over the range of the system. Note that some transducers can exhibit hysteresis from being exercised fully, so always allow 5-10 minutes for the transducers to settle after performing an auto exercise to start testing.

What is the proper warm-up time for an ADTS?

Temperature sensitivity is minimized on the ADTS-3350ER through the use of a temperature monitoring system, that measures the temperature in several areas internally. The ADTS can be ready to use in a few minutes under normal conditions.  On a really cold day, it may take the

 entire 10-15 minutes for the internal temperature to stabilize.

Can you control the airspeed without the static hose being attached to the aircraft?

You really shouldn’t.  If all you want to check is your indicated airspeed, you should still connect the static hose to the aircraft. The reason for this is that the ADTS altitude may not always be exactly at ambient pressure.  If this is the case, your airspeed readings will be a little off.

Also, if for some reason, the ADTS altitude is increased with only the pitot hose attached, you will generate a negative airspeed into the airspeed indicator, Mach indicator, air data computer, or any other differential pressure-sensing instrument connected to the system.

For example, let’s say the outside pressure is 30.000 inHg.  The ADTS is connected to the pitot system only with 200 knots applied to the system.  (Remember, at sea level, 200 knots is equal to about 31.900 inHg.  At 10,000 feet, 200 knots is only about 22.500 inHg.)

If, for whatever reason, the ADTS altitude is increased, the ADTS will automatically decrease the pitot pressure in the lines to maintain 200 knots for whatever altitude is on the test set.  Meanwhile, you still have the 30 inHg inside the static system causing the airspeed indicator to drive off scale low.  You will most certainly damage any instrument that is connected to the pitot system if you do this.

Remember that you must always maintain a pitot pressure that is equal to or greater than your static pressure.  The easiest way to do this is to always use both hoses for whatever you do.

The post Air Data Principles and Aircraft Pitot-Static System Troubleshooting using the ADTS-3350ER or ADTS-3300JS Air Data Test Sets appeared first on Raptor Scientific.

(no names mentioned)

This post contains a description of actual case histories of instances where a mass properties measurement error has occurred which could have been avoided (no names mentioned!). Some of these errors resulted from fundamental defects in the procedure; others resulted from very subtle effects which would have been hard to anticipate. And others are painfully obvious, once you realize the problem. However, the fact that these errors occurred emphasizes the need to have a checklist, and to have a person who is knowledgeable in mass properties supervising the measurement.

The purpose in publishing this history of real errors is to help you avoid these pitfalls. Many of us have made these mistakes (sometimes more than once). We hope that persons reading this paper will supply additional examples to us, so that we can publish a sequel to this in the future.

Case 1: A Multiple Scale System Which Didn’t Work

A company did not have a single scale that was large enough to weigh a vehicle, so the solution was to place three scales under a platform to create a larger scale. The vehicle was placed on the platform, and then the three scale readouts were added. When this method had been devised, it was hoped that a crude measurement of center of gravity could be obtained from the difference in the scale readings (if all three scales read the same amount, the cg was midway between the scales, etc). However, when this measurement was attempted, it was impossible to get repeatable readings. The combined weight would change when the vehicle was moved from one place to another on the platform, and the measured cg did not agree with the distance the vehicle was moved on the platform. In other words, if the vehicle was moved by 1 inch in the X direction, the calculated cg did not change by 1 inch.

The first step in identifying the problem was to have the three scales calibrated. They were sent to the lab, and they all were shown to have errors that were less than 0.05%, yet the sum of the three readings had varied by more than 1%!

The second step in identifying the problem was to re-assemble the platform and place large test weights on it. The readings changed when the weights were moved. In addition, the total measured weight was larger than it should be. What was wrong??

{Answer}: Scales are not made to be linked together. In order to get accurate readings, a scale must float freely and not have any sideload. Some scales are more sensitive to this problem than others. In this particular case there was another problem: the thin platform was bending under the weight of the vehicle, so that it was not square with the top of the scale, and presented a different side load when the vehicle was moved from one position to another. (see appendix for more info)

Lesson Learned: Scales cannot be linked together. They are only accurate when operated independently.

CASE #2 Multiple Load Cell CG Problems

The next attempt in measuring the same vehicle was to hang the vehicle platform from three load cells. Since the roof of the building was not strong enough to support the weight of the vehicle, a structure had to be constructed. The three load cells were attached to this structure, and the platform was hung from three cables. Flexures were used on either end of the load cells to make sure there was no bending moment(this organization was not going to make the same mistake twice). When the vehicle was measured, the measured weight was constant, no matter what position the vehicle was placed on the platform. Success! But wait a minute: when the vehicle was moved 1 inch, the calculated cg position moved much more than 1 inch. Now, what was wrong? How could the cg be in error when the total weight appeared to be correct?

{Answer}: The cables were stretching, causing the vehicle to lean. Since the vehicle cg was considerably above the platform, the cg would move outward when the cable stretched, amplifying the cg offset. This was determined by placing a level on the platform, and then moving the vehicle to a new position.

Lesson learned: Multiple-point cg measurements require a rigid force measurement system. Otherwise, the cg of the object will lean outward, amplifying the cg offset.

CASE #3 The Levitating Satellite which Turned Cartwheels

It is essential that the center of gravity of a satellite be exactly in line with the centerline of the thrusters. Otherwise, when the thruster is fired, a torque will be created which causes the satellite to spin rather than translate its position. In this particular tragic case, a satellite test platform had been constructed to demonstrate the operation of a new control concept. In this experiment, the test satellite was to be dropped above a net, and the thrusters were to be fired while the satellite was in midair. It was expected that the satellite would be suspended in air and could be maneuvered above the net for a number of seconds. High-level government personnel were present to witness the test. The day prior to making the test, the center of gravity had been measured and correction weights were added to move the CG to the exact centerline. Unfortunately, no one thought to re-measure the cg after the weights had been added. They had been added to the wrong side, doubling the cg offset, rather than trimming it to zero. At the appointed moment, the high-speed video cameras were started, and the satellites were dropped. Instead of levitating above the net, it instantly went into a spin.

Lesson learned: Always re-measure a vehicle after weight correction has been performed to make sure that the correction was successful.

CASE #4 Turbine Blade Fixturing Error

Two facilities were measuring the moment weight of the same type of blade. There was a consistent difference between the measurements made by the two facilities so that blades measured by one could not be mixed with blades measured by the other when assembling the jet engine rotor. It turned out that the adapter had been assembled differently at the two facilities. At one location, the adapter spring caused the turbine blade to be forced forward. At the other facility, the adapter forced the blade backward. Which one was correct?

{Answer} Centrifugal force pulls the blade forward when it is spinning in the engine. Therefore, the blade should be forced forward in the adapter. The design used by company B is correct.

CASE #5 Another Turbine Blade Fixturing Error

Turbine blades in a jet engine must be sorted by cg moment. Pairs of blades with identical moment are installed in the engine 180 degrees from each other, so that the unbalance is minimized. A jet engine manufacturer was sorting these blades with what he thought was a high precision, yet the engines were badly unbalanced after the blades were installed.

The problem was the point of contact of the blades with the adapter that was used in the moment measuring scale. In an engine, centrifugal force pulls the blades outward until the root of the blade contacts the hub of the engine. The engine manufacturer was contacting these blades at a different point during his static moment measurement.

CASE #6 The Case of the Errant Allen Wrench

A company did a statistical analysis on the center of gravity data for a particular object, and they discovered that the first shift operator consistently got an X-axis cg location which was about 1/8 inch different from the second shift operator. To verify this observation, they had the first shift operator and the second shift operator both measure the same object. Sure enough, this measurement was also 1/8 inch different. An engineer was assigned to determine the cause of this difference. The illustrations on the next page give a clue to one part of the problem. As you can see, operator #2 put the Allen wrench on his workbench after tightening the screws which mounted the object to the fixture. Operator #1 put the Allen wrench in a hole in the fixture. This was a handy place to keep the Allen wrench, but unfortunately, it added an unbalance mass to the measurement. Since the tare CG had been determined without the Allen wrench, this mass caused a consistent cg error on his measurements.

At first, it appeared that the problem had been solved. However, after the correct procedure was established for storing the Allen wrench, the cg data was still different between operator #1 and operator #2. What was going on? The illustration below of the configuration during tare measurement gives a clue to the answer:

{Answer}: Operator #2 included the mounting screws in the tare measurement because he reasoned that these screws were not part of the object, and therefore should be left in place during tare measurement so their contribution to unbalance moment would be subtracted from the object measurement. Operator #1 left these screws out during tare measurement. Although operator #1 was careless about his location for the Allen wrench, he made the right decision about the mounting screws. These are installed during flight and therefore constitute part of the object.

Lesson learned: You must make a decision whether the mounting hardware is part of the object or unique to the fixturing method. If it is part of the object, then it should be removed during tare measurement. Don’t leave any loose items such as wrenches on the fixture during measurement.

CASE #7 Difference in MOI between Measurement on Earth and Flight in Space

The moment of inertia of a satellite was measured in a lab prior to flight. After the satellite was placed in orbit, the response to thruster correction seemed to indicate that the MOI of the satellite was smaller than measured on earth. What was the reason?

{Answer}: For large lightweight payloads, the measured mass properties in air are often significantly different from the values in the vacuum of space. In particular, measured moment of inertia can be 10% to 20% larger than calculated. The reason for this is that air has significant mass and alters the mass properties in two ways:

• Air trapped inside the payload will increase its mass by an amount equal to the unoccupied volume in the payload times the density of air (0.0754 pounds per cubic foot). For example, the air trapped in a 6-foot diameter satellite might weigh approximately 4 lbs. We call this the entrapped air effect.
• Air dragged or pushed along by any protrusions on the outer surface of the payload can dramatically increase moment of inertia. For example, the roll moment of inertia of an aircraft flying at sea level is larger than the roll MOI of the aircraft at 40,000 feet. We call this the entrained air effect.

If the payload flies in vacuum, then measured values must be decreased to eliminate the effect of air mass. The best way of doing this is to make a second  measurement in helium and then extrapolate the value in vacuum (see SAWEpaper No. 2024 by Boynton and Wiener). Or if you have a vacuum chamber which
is large enough to accommodate the satellite and mass properties machine, you can measure the satellite in a vacuum. However, this requires special modifications to the mass properties machine.

CASE #8 Does RoomSize Affect Measurement Accuracy?

The moment of inertia of an airfoil was measured in a small experimental lab. Readings were very repeatable. However, when the machine was moved to the large production floor with a 60-foot high ceiling, it was impossible to get consistent readings. This area had an overhead door which led to the outside of the building, but the door was always closed before measurements were made, and the air conditioner was shut off to prevent drafts. Why was repeatability worse in a larger room?

{Answer}: The problem turned out to be air inversion. If the overhead door was opened for even a short interval, warm air entered the building at the ground level. Meanwhile, cool air was leaving the A/C duct at the ceiling of the building. This created an inversion layer. The warm air continued to rise long after the overhead door was shut, creating drafts which acted on the airfoil section.

CASE #9 Importance of Axis Definition

The cg of a 22 inch long projectile was measured at one facility. The cg machine and the fixture were then shipped to another facility and the measurement was repeated. By analyzing the data, see if you can figure out why the answers were different.

{Answer}: The nose of the projectile contacted the fixture end stop in measurement #1, whereas the aft end of the projectile contacted the fixture end stop in measurement #2. When the cg measurement specification was reviewed, it turned out that it did not define which end was the reference. Therefore, each  measurement technician inserted the projectile in the direction that seemed logical to him.

Lesson learned: Measurement specifications must define the axes! Use the SAWE Recommended Practice #6 Standard Coordinate System whenever possible.

CASE #10 Misjudging the Effect of Electrical Cable Weight

We sold a gimbal balance machine to a company overseas. This instrument measures the CG of a seeker with extraordinary precision. We can detect an unbalance of 0.0001 lb-inch. Soon after we sold the gimbal balance machine, we learned that the customer was having a problem with repeatability. I visited this customer and discovered that he had attached a 1/4 inch diameter cable to the part he was measuring. The cable was hanging over the side of the machine, producing a moment that was about 1000 times what he was allowed. He had recognized this problem and had added a weight to the gimbal to compensate. What he didn’t realize was that a minute motion of the cable would produce a change in moment that was at least 10 times his tolerance.

Several years later, the problem reappeared. He had learned to avoid attaching anything to the gimbal and was mystified that the repeatability problem was back. The difficulty turned out to be the cables inside a new type of gimbal. Numerous heavy cables had been run from the rotating assembly in the gimbal to the base of the gimbal. Therefore, the weight of the cables contributed to unbalance. A very small change in the position of these cables upset the balance. It was necessary to redesign the gimbal to eliminate this problem.

CASE #11 Protective Paper Usedonthe Face of a Gimbaled Seeker

Another company purchased a gimbal balance machine from Raptor Scientific for the measurement of a gimbaled seeker. This seeker had a protective paper cover to prevent damage to the microwave antenna assembly. Balancing was performed with this paper in place. How much error could this paper have caused?

{Answer} Enough to make the seeker unbalance 10 times greater than the unbalance allowed.

CASE #12 Using Balancing Clay to Predict the Ballast Required

This is a problem that we have seen many times. A vehicle is placed on a spin balance machine, and balancing clay is added to several locations on the vehicle skin until a balance is achieved. The position of these lumps of clay is then determined, and they are removed and weighed. Ballast weight are then fabricated. Their weight is identical to the lumps of balancing clay. The outer covers of the vehicle are then removed, and these weights are installed inside the vehicle. The vehicle covers are then re-installed, and a final spin balance measurement is performed. The measurement indicates that the weights were too light. Why?

Answer}:{ The radius of the balancing clay is greater than the radius at which the ballast weight are installed. Unbalance is proportional to the moment RW, where R is the radius to the CG of the ballast weight (or the CG of the balancing clay). Let’s say that the initial unbalance was 100 lb-in². Normally you might expect the unbalance after correction to be about 3 lb-in². However, if the radius to the CG of the balancing clay was 12 inches, and the radius to the CG of the ballast weight was 9 inches, then the unbalance after correction will be 25 lb-in², rather than 3 lb-in². To avoid this problem, you must increase the ballast weight by the ratio of the balancing clay Cg radius divided by the ballast CG radius.

Lesson learned on case #11 and Case #12: Make certain that non-flight items are removed before weighing.

Errors that No One would be Dumb Enough to Do (but they did)

1. A common method of achieving redundancy is to put two load cells in series. Since each is measuring the same force, they should agree within a small percentage. One mass properties engineer sent a work order to have the weight of several large objects measured, using a crane hook scale. The mass properties engineer was not given the authorization to witness the measurement. When the data was sent to him, the values were almost exactly twice what he had calculated. What was wrong?

{Answer}: The test technician had added the readings from both load cells to get the total weight.

2. An accurate fixture is required to position a test object at a precisely known position relative to the mounting plate of a  cg instrument. Fixturing inaccuracy is usually the major source of cg measurement error. However, occasionally we encounter an engineer who doesn’t even realize that a fixture is necessary. Someone called Raptor Scientific a number
of years ago, after receiving a cg instrument. His complaint: “I can get any answer I want from your machine. It depends where I place my object on the cg instrument mounting plate. If I move the object by 1 inch, then the answer changes by 1 inch.”

Lesson learned: Even if their rate of pay is low, it doesn’t pay to use unqualified personnel.

Appendix

Why scales cannot be linked together mechanically without introducing error

Weight scales are designed to be free-floating. The internal structure consists of a parallelogram, which causes the scale to follow a specific deflection path when weight is added to the scale. If the scale is leveled properly, this path is approximately (but not exactly) vertical. When two or three scales are linked by a platform, the application of weight causes the scales to be forced apart or drawn together, depending on the misalignment between the scales. Furthermore, the platform usually deflects, introducing another component of misalignment.

Multiple Point Weighing Method

To determine part weight (W) and CG coordinates X and Y, three force transducers are typically used to support a frame which in turn supports the test part.

W = A + B + C where A, B, and C are force readings on the three force transducers.

To determine CG, take moments about A

where X and Y are the CG coordinates. If all three scale outputs are set to zero when fixturing is in place, the equations above cab be used to determine the CG  location of the test part. In practice, tare readings are subtracted from the part.

The problem is that even a small error due to side load translates to a big CG error. Assume that

• D = 60 inch
• Side load error = 2% of weight W

Then

• Y = 0.5 x 60 x 0.02 = 0.060 inch

Note: Aircraft weighing generally involves three-point measurement. There are two reasons why side load errors are minimized in this process:

1. The scales which are used are specially designed to minimize side load errors.
2. As the aircraft is rolled onto the weight platforms, the rolling action tends to minimize the sideload. In some instances, aircraft are rolled back and forth to further reduce this effect. This situation differs from the use of a solid platform which is coupled directly to the three scales.

About the Author

Richard Boynton was the President of Space Electronics, Inc., Berlin, Connecticut, a company he founded in 1959, and which was purchased by Raptor Scientific. Mr. Boynton has designed many of the mass properties measuring instruments manufactured by Raptor Scientific. He holds a B.E. degree in Electrical Engineering from Yale University and has completed graduate studies in Mechanical Engineering at Yale and M.I.T. He is the author or co-author of over 69 papers, including 30 papers presented at SAWE International Conferences and 3 papers presented at regional conferences. He is the author of the SAWE Recommended Practice for Standard Coordinate Systems for Reporting the Mass Properties of Flight Vehicles. Mr. Boynton has been a member of SAWE for over 30 years and is currently Director of the Boston Chapter. In 1992 he was elected a Fellow and in1998 was elected an honorary fellow of the SAWE. Mr. Boynton is a member of the Society of Automotive Engineers, where he serves on the Balancing Subcommittee (which is currently involved with setting standards for jet engine balancing).

The post Mass Properties Measurement Errors Which Could Have Been Easily Avoided appeared first on Raptor Scientific.

Measuring the resistance of an igniter used to detonate an explosive is a little like measuring the length of a live rattlesnake . . . the measurement itself is relatively simple, but there are certain safety aspects that strongly affect the design of the measuring apparatus. In plain English, when you test an igniter circuit, you will die if the test current emitted from the resistance measuring instrument accidentally exceeds “safe” limits. For most electrical instruments the design engineer spends 99% of his time worrying about measurement accuracy and performance features, and then almost as an afterthought, he briefly considers the safety aspects of the tester and the consequences of a malfunction in the circuitry. When designing an igniter circuit tester, the emphasis is just the opposite. Thousands of hours are spent thinking about the safety of the tester. If there is a 1 in 1,000,000 chance that some freak combination of malfunctions and operator errors could cause excess test current, sending an armed ICBM toward New York City, then safeguards must be built into the tester to prevent this condition.

What is an Igniter?

Modern rockets are “set off” by passing a current through a small heating element imbedded in inflammable material. When the temperature of this small heating element reaches a critical value, the “igniter” bursts into flame, initiating the rocket fuel. This same process is used to set off weapons–sometimes the igniter creates a small explosion which triggers the main explosive material; other times the igniter burns, and the resulting gas pressure operates a small cylinder which begins a sequence of mechanical and/or electrical steps leading to the detonation of the warhead. A third process involves the use of “explosive bolts”–fasteners that instantaneously fly apart when a current is passed through the small heating element in the igniter within the bolt. These same bolts are used extensively in space vehicles and underwater devices to rapidly separate large objects without requiring heavy, large mechanisms.

How is an Igniter Tested?

An igniter circuit tester is a special type of ohmmeter that measures low resistance very accurately, using a test current that is as much as 1000 times less than a conventional precision low resistance ohmmeter. The low test current is necessary to prevent ignition of the rocket motor or detonation of the explosive device being tested. Since excess test current could result in bodily injury or death, every possible way of accidentally creating overcurrent must be considered and protected against. This kind of thinking leads to such tests as deliberately failing components to see what will happen (certain resistors must fail open rather than shorted, for example). Quality assurance testing must exceed even the tightest “MIL Spec” procedures, and finally, the designer must be able to second-guess all the mistakes any test operator might make in using the tester. It must not be possible to connect the leads to the tester improperly, and it must be virtually impossible for someone to open the instrument case and modify or bypass the safety circuit elements (which prevent overcurrent in the event of a failure of the active circuits in the tester). All critical elements must be redundant and must fail in a safe way, causing test current to be less than 5% of the value required for ignition or detonation of the part EVEN IF ALL ACTIVE ELEMENTS OF THE CIRCUIT FAIL SIMULTANEOUSLY.

In the many years we have been in the business of testing explosive devices we think we have seen most of what can go wrong–batteries leaking and shorting out circuitry, technicians opening the case of the instrument and bypassing the current limiting devices, inexperienced personnel replacing the internal battery with one of higher voltage or, God forbid, wiring in an AC supply to replace the batteries without realizing that this supply must have an “ultra isolation” shielded transformer and zener fail-safe networks to prevent potential differences between earth ground and AC ground from detonating the part.

Our latest instruments have every safeguard we can think of. Critical fail-safe circuitry is sealed in a potted assembly attached directly to the output connector. The AC charger must be unplugged in order to connect the test leads, making it impossible for a test to be run with the AC still connected. We use a special rechargeable nickel-cadmium battery pack which is totally sealed in a thick plastic case to guard against leakage and has an output connector that does not match any off-the-shelf battery of higher voltage. Instruments are subject to a series of environmental and performance tests, ensuring both accuracy and reliability.

To our knowledge, our instruments are the only automatic igniter circuit testers approved by the military as safe for testing explosive devices.

The Accuracy Problem

It is very important for an igniter circuit tester to be accurate: certain problems such as a weak spring on a safety shorting switch manifest themselves in resistance changes as small as 0.1 ohm. Typical accuracy of a high-quality igniter circuit tester is 0.010 ohm with resolution as high as 0.001 ohm. This kind of accuracy is readily obtainable with a precision lab Kelvin bridge–you just pump enough current through the unknown resistance to get a big voltage drop. The magnitude of current used in a typical lab bridge would set off any igniter. In igniter circuit testers, accuracy must be obtained with voltage drops as small as 10 microvolts. Such requirements are right at the state of the art in low-noise low-drift balanced differential operational amplifiers. In some systems, obscure effects such as thermoelectric voltages become the predominant limitations to accuracy.

Safety Warning

There are some instruments on the market that can explode the device being tested if the instrument malfunctions. Other instruments permit fail-safe circuitry to be bypassed. Be wary of “lab type” instruments that have been modified for igniter circuit testing. Testing explosive devices can be a matter of life and death.

Beware of the Old Wheatstone-Bridge-Type Testers

In the 1950’s, most igniter circuit testers consisted of a Wheatstone bridge with a current limiting resistor in series with the internal battery. Measurements were tedious with this type of tester (you had to null the bridge for each measurement and then subtract the test lead resistance–these were two-wire type bridges); but most important, THESE INSTRUMENTS WERE NOT VERY SAFE. On these old testers, a technician could easily open the instrument case and modify the circuit, inadvertently bypassing the current limiting resistor. This has resulted in detonation of an explosive device in a number of occasions over the last 20 years. Unfortunately, many of these old instruments are still in use. They present one particularly dangerous hazard–if a replacement battery is not available (the old testers did not have rechargeable batteries), technicians sometimes wire a different battery into the unit. Since the current limiting resistor on these old testers was attached to the battery case, the bare exposed terminal on the outside of the case is

“downstream” of the safety circuitry. Almost anyone looking at the inside of the tester would think this terminal is the output from the battery; but, in fact, if a battery is connected to this point, the igniter circuit tester will output over 100 MA, detonating most parts. This tester had other problems. Inexperienced users could misunderstand the operation of the instrument and fail to null the meter on the correct range. Or they would forget to subtract test lead resistance (Space Electronics testers do this automatically using a true four-wire “Kelvin” circuit).

Igniter Circuit Tests

There are two basic tests that are performed on all igniter circuits:

A. Continuity. Most igniters have a resistance of about 1 ohm. After the igniter is installed in its final configuration, it is necessary to verify that all the associated circuits (including the igniter itself) are functional. Obviously, the system cannot be operated to confirm its performance (press the button and if the bomb goes off, everything was working). The best test is to measure the total resistance of the igniter circuitry. A very accurate ohmmeter is required, since a short might only change the total resistance from 1.8 ohm (for example) to 0.8 ohm. Even more critical is the contact resistance of the switch elements in the circuit–a 0.2 ohm change can often give a clue to a damaged contact spring or improperly installed mechanism.

There are actually at least two low resistance measurements made on an igniter circuit. In addition to the series resistance of the igniter element, it is also necessary to measure the resistance of the safety shorting switch which is part of the design of almost every rocket igniter or warhead detonator. This shorting switch is in parallel with the igniter; a variety of conditions must exist before this switch is opened, allowing current to flow through the igniter (for example, a projectile must experience acceleration and then deceleration before the switch opens on the warhead igniter). Since this switch is in parallel with the igniter, test current must be safe limited–in the event the switch were accidentally open (the condition being tested for), excess test current would set off the explosive.

B. Open Circuit Test. In addition to continuity checks, it is essential to verify that the igniter is NOT inadvertently connected to some other circuit, either by virtue of a wiring error, or by the igniter lead wire being pinched under a mounting bracket or any of the thousands of other mishaps that can occur in, for example, the circuit that fires the explosive separation bolts in a two-stage rocket. This test generally has a MINIMUM acceptable resistance between 100K and 2 Megohms. Usually this test involves checking each input to the igniter circuit relative to the ground connection or case of the device.

Safety Guidelines

Although the safety of igniter circuit testers is extraordinarily high, nothing is perfect in this world and there are some common-sense rules that I would recommend:

1. If the part can kill, test it behind a barrier even if you are certain the test is 100% safe.
2. If you can’t see the tester while you are installing the leads on the test part, use a shorting switch that connects all the test leads together while you are working near the part (Space Electronics makes a key-operated switch attached to the test cable; you can take the key with you and not worry about who is fooling around with the tester buttons while
3. you are connecting leads to a bomb that can blow a 15-foot hole in the ground).
4. Don’t ground anything in the circuit; or if you must, ground only one point. Earth ground can differ from AC ground by as much as 5 volts: this can result in a current of amps in some instances (Our testers are battery-operated, so this problem isn’t a safety consideration).
5. Always check “Open Circuit” as well as continuity, whether the test is required or not. It’s your life that is at stake.
6. Do 100% inspection.
7. Test again if ANY modifications are made to the part or if any part is dropped or jarred. If the part is disassembled and put back together, test it again.
8. Wherever possible, connect the tester to the part using a mating connector that has been carefully wired and inspected. Alligator clips can give false readings if they accidentally short against the case–I have seen instances where a defective part passed a test only because the test clips shorted in such a way to result in a resistance that happened to fall within acceptable limits.
9. Check any explosive device at every stage of production if you can, but at least check the entire arming circuit with the device unloaded, before making the final check with the full explosive installed. Aside from safety considerations, this procedure greatly simplifies the disposition of the part if it fails the test.
10. If a test fails, evacuate the area, and go through the full disarmament routine. Don’t wander behind the barrier to see if the test leads are tight (you should have done this when you installed them).
11. Test small groups of live explosives at one time. Then move this group out of the area before starting another series of tests. I have seen test areas where two weeks’ worth of explosives were piled to the ceiling because the usual material handler had been sick.
12. If the test part leads are not color-coded according to the drawing or the way they were last week, stop the test and get positive clarification. Don’t assume a new vendor uses other color codes or that the tester is so fail-safe that you can experiment with the part. The new vendor may have shipped you 5,000 warheads designed with an igniter whose safe current is 40 microamps, and you may have an igniter circuit tester that puts out 5 milliamps. Every one of the warheads will detonate when you test it.
13. Don’t test during lightning storms.
14. If you are using an AC power pack option with your tester (not recommended for high-risk parts), measure ground potential differences between AC ground pin and earth ground before doing any testing. If it’s more than 0.1 volt, then something is wrong

The post Igniter Circuit Testing appeared first on Raptor Scientific.

Secret Number 1 – The payload must have precisely defined measurement axes. Both your calculated and your measured data are only as good as your ability to define the axes of the payload.

Secret Number 2 – The fixture must hold the payload so its measurement axes are precisely located relative to the instrument. Fixturing error is the number one source of measurement error in most mass properties measurements.

Secret Number 3 – You must follow the correct measurement procedure. The four most common procedural errors are:

1. Tare measurement is made with mounting bolts left out or clamp set in a different position than when the measurement was
2. Payload is not in flight configuration (arming switch must be set, shipping clamps must be removed, protective covers must be removed, tanks must be filled with fuel).
3. Fixture unbalance is too large. This must be smaller than the desired balance. Otherwise you will be subtracting two large numbers. A small error in either will result in a large error in the difference. Also, the machine will have to be set on a low sensitivity
4. Payload is fixtured so its CG is too far from balancing machine axis. (This is similar to the problem described above). This also results in overturning moment stiffness Since no balancing machine is infinitely stiff to an overturning moment due to CG offset, all payloads lean toward their CG. To get the most accurate measurements, reposition the payload so its CG is on center and remeasure.

Secret Number 4 – You must eliminate external influences (drafts, temperature changes, vibration, air mass).

Secret Number 5 – The payload must weigh more than 2% of the capacity of the mass properties machine. Instrument accuracy is generally a function of payload weight or moment of inertia. If you try to measure a small object on a large machine, then it is like trying to weigh yourself on a 20, 000-pound truck scale; the answer gets lost in the background noise.

Secret Number 6 – You must use an accurate instrument. Sensitivity and repeatability are not the same as accuracy. If your calibration standard is in error, then all of your measurements will also be in error. If the payload leans away from the centerline of the instrument, this will result in a large measurement error.

Secret Number 7 – You must define your symbols and polarities.  Define which axis is X, which axis is Y, etc. Your X may be someone else’s Y. Even within one company, one department may call the roll axis X and another department may call it Y. If you submit the data without defining the axes, each group will use its own set of coordinates in interpreting the data.

Secret Number 1. The payload must have precisely defined measurement axes.

The mass properties of an object cannot be separated from the axes used as a reference. In fact any object has an infinite number of values for CG, moment of inertia, and product of inertia, depending on where you choose to assign the axes. If the object is a smooth ground cylinder, then it is obvious where the axes are located. However, on real parts,

• flat surfaces are not perfectly flat;
• round surfaces are not perfectly round;
• concentric surfaces are not exactly on the same center;
• perpendicular surfaces are not exactly perpendicular;
• some surfaces are soft (cork, thick paint).

STEP 1 = Calculate the required mechanical dimensional tolerances necessary in order to be 10 times better than the accuracy specification for mass properties. For example, if CG accuracy required is 0.005 inch, then you must know the location of the reference axes to an accuracy better than 0.0005 inch. Tolerances for product of inertia are trickier to calculate. The best approach is to first calculate the axis tilt corresponding to the POI tolerance, and then relate this to TIR runout of two reference diameters.

STEP 2 = Do a dimensional inspection of the payload. If the payload outer surface is less accurate than the tolerances calculated in step 1, then you are in trouble.

NOTE: Raptor Scientific provides a measurement service to small companies who do not have the funds to purchase a mass properties instrument. For at least 30% of the measurements we are asked to do, we discover that the dimensional tolerances of the object are not tight enough to permit the mass properties accuracy required!

QUESTION: If the dimensional tolerances are not sufficient to achieve the accuracy required, what do you do?

ANSWER: You must know why the specification was important, and then you can devise some means of accomplishing the real goal. For example,  the real reason for controlling CG may be to put the CG on the center of thrust. If you can locate the CG relative to the centerline of the nozzle, then you can accomplish this goal even if the outer surface of the vehicle is made of soft cork. This topic is discussed in some detail in the SAWE paper number 2101 entitled “An Expanded Role for the Mass Properties Engineer” by Richard Boynton, President, Space Electronics, Inc.

QUESTION: What do you do if the mass properties measurement test plan requires you to measure CG relative to a nebulous datum, when you know the purpose of this measurement is to align the CG with the center of thrust?

ANSWER:

1. Do the right thing. Measure the CG relative to the center of thrust and make the proper correction.
2. Fudge the data to keep the inspector happy (i.e. after you do the right thing, find a way of expressing the data so you pass the test).
3. Try to get the test changed so you won’t have to play games in the future.

If you have an influence in the early stages of a design, maybe you can convince the project engineer to add two precision datum rings to the payload. This will give you a reliable interface for your fixture and will also give you something to measure to determine if the payload is located correctly in the fixture. Engineers who align the guidance system will find these rings invaluable. Thruster nozzles can be located relative to these rings.

Secret Number 2. The fixture must hold the payload so its measurement axes are precisely located relative to the instrument.

Fixturing error is the biggest error in most measurements. Years ago a customer called us after receiving a mass properties instrument and said “There’ s something wrong with the instrument; I can get any answer I want by just moving the payload to a new position on the mounting plate of the instrument.” We all got a good laugh at Space Electronics, and frankly we thought the customer was a real moron.  Since then we have come to realize that many people do not fully realize the importance of accurate fixturing.

POI: Even the smallest tilt is very significant; payloads should be dial indicated whenever possible. Do not rely on the fit in a fixture to establish position.

CG: A 0.001-inch fixturing error translates to a 0.001-inch measurement error.  It is very difficult to get this kind of accuracy unless the payload has a precision outer surface.

MOI: Fixturing accuracy is not critical except if the payload is tall and thin.  The reason for this is that the error is proportional to the square of the ratio of radius of gyration (“k”) and fixture offset error (” d”). Generally the fixturing error is less than 1% of the radius of gyration, so the resulting error will be less than 0. 01%. This relationship is derived below, using the well-known formula for translations of axes.

Using a Precision Dummy Payload One very convincing method to verify fixture accuracy is to construct a precision test weight with known mass properties which interfaces with the fixture in the same way as the real payload. For example, this weight might be a simple cylinder of constant diameter. If the mass of a solid cylinder would be too large, but you need a large diameter to interface with the fixture, you can use a small diameter solid cylinder with a larger diameter disc attached to each end.

Work reversal method (only works for symmetrical objects) Errors in fixture position relative to the measurement axis of the instrument can be eliminated by turning the object 180o in the fixture and remeasuring unbalance. The unbalance magnitude should stay the same; angle should change by 180o. If you average the results and divide by two, then fixture position error is washed out.

Note: this method eliminates fixture position error, but does not eliminate errors due to poor fit in the fixture.

Secret Number 3. You must follow the correct measurement procedure.

This means you must have a written procedure! Write it as a step-by-step guide. Don’ t say “Measure tare”. Instead say ” Loosen the clamp screw, remove the payload from the fixture and then tighten the clamp screw so it is in exactly the same position as when the payload was in place. etc”.

Set limits in the procedure. Don’t say “Center payload in fixture”. Say “Center payload in fixture so upper and lower rings run out less than 0. 001 TIR”.

Be very careful how to describe how the payload fits in the fixture. This is particularly important when the payload can be accidentally put in backwards. Something which may be obvious to you will not necessarily be obvious to someone else.

The procedure should caution the technician about burrs, dirt between payload and fixture, misalignment.

The procedure must be specific for the payload being measured (other payloads will have to positioned differently, even if the same fixture is being used).

The four most common procedural errors:

1. Tare measurement is made with mounting bolts left out or clamp set in a different position than when the measurement was made.

2. Payload is not in flight configuration (arming switch must be set, shipping clamps must be removed, protective covers must be removed, tanks must be filled with fuel).

3. Fixture unbalance is too large. This must be smaller than the desired balance. Otherwise you will be subtracting two large numbers. A small error is either will result in a large error in the difference. Also, the machine will have to be set on a low sensitivity range.

4. Payload is fixtured so its CG is too far from balancing machine axis. (This is similar to the problem described above). This also results in overturning moment stiffness errors. Since no balancing machine is infinitely stiff to an overturning moment due to CG offset, all payloads lean toward their CG. To get the most accurate measurements, reposition the payload so its CG is on center and remeasure.

Secret Number 4. Eliminate external influences (drafts, temperature changes, vibration, air mass).

Any external force will degrade the accuracy of a mass properties machine. However, each kind of measurement is more sensitive to a certain kind of external influence.

CG is most sensitive to temperature changes. Generally we recommend that the short term changes be less than 1o F per 15 minutes. This does not require an expensive air conditioner. A simple shroud will generally do the trick. Make certain the output of the heating/air conditioning system is not near the instrument.

MOI is most sensitive to drafts. This will be immediately obvious, since the drafts will cause the time period to show random variations. In the proper environment, time period reading should be repeatable within three parts in 100,000.

POI is most sensitive to ground vibration. Static CG machines filter out vibration forces, since the average of these forces is zero. However, a spin balance machine measures the sinusoidal vibration at the bearing mounts, so the force transducers have to be able to detect vibration. Tracking filters in the machine eliminate vibration whose frequency does not coincide with the rotational speed, but any external signal whose frequency is close to the rotational speed of the machine will cause an error in measurement. Do not locate a spin balance machine near air compressors, “shakers” used in vibration analysis and testing, or lift truck traffic.

Effect of air mass For large lightweight payloads, the measured mass properties are often significantly different from the calculated values. In particular, measured moment of inertia can be 10% to 20% larger than calculated. The reason for this is that air has significant mass and alters the mass properties in two ways:

Air trapped inside the payload will increase its mass by an amount equal to the unoccupied volume in the payload times the density of air (0.0754 pounds per cubic foot). For example, the air trapped in a 6 foot diameter satellite might weigh approximately 4 lbs. We call this the entrapped air effect.

Air dragged or pushed along by any protrusions on the outer surface of the payload can dramatically increase moment of inertia. For example, the roll moment of inertia of an aircraft flying in air is much larger than the roll MOI of the aircraft in a vacuum. We call this the entrained air effect.

How you handle this difference depends on whether the payload operates in the vacuum of space or in air. If the payload flies in a vacuum, then measured values must be decreased to eliminate the effect of air mass. The best way of doing this is to make a second measurement in helium and then extrapolate the value in vacuum (see SAWE paper No. 2024 by Boynton and Wiener). Calculated values remain unchanged.

If the payload flies in air, then measured values remain unchanged and represent the true mass properties. Calculated values should be changed to reflect the effect of air mass.

Secret Number 5. Buy an accurate instrument.

The instrument must be at least 5 times more accurate than the required tolerance on mass properties.

Accuracy versus sensitivity A CG sensitivity of 0. 001 inch tells you that if you move the payload by 0.001 inch, the machine will be able to detect that you made this change. A CG accuracy of 1% means that if you measure a CG offset of 0.100 inch, then the true offset is between 0. 099″ and 0.101″. There is a big difference between these two specifications.

Machine measurement axis Machines that use rotating mounting plates have a clearly defined axis (the center of rotation). The payload can be very accurately aligned to this axis by using a dial indicator. Load cell tables may be sensitive to CG change, but there is no well defined zero or measurement axis. Machines with gas bearing rotary tables are the most accurate.

Calibration weight errors Since the “scale factor” of a mass properties instrument is determined by measuring a known calibration weight, an error in the calibration magnitude will result in consistent errors in all measurements. Calibration weights should be precision machined and be traceable to NIST.

Stiffness to overturning moment A CG instrument may be accurate when measuring low profile payloads, but the accuracy can degrade very rapidly if the payload is tall and thin. The reason for this is the displacement of the payload due to lean. This problem is particularly evident on three-load-cell type CG instruments which hang the test table from aircraft cable. If the payload CG is displaced from the center of the machine, then there is a moment created. This moment causes the table to lean toward the CG, increasing the apparent CG offset. The higher the CG from the surface of the table, the more pronounced this effect will be. Instruments which use rebalance technology (such as the KSR series) do not exhibit this type of error because the closed loop feedback keeps the table level independent of CG offset.

INSTRUMENT ACCURACY TESTS:

1. Repeatability of instrument (Leave object in fixture and )
2. Repeatability of fixture (Remove object from fixture and replace; then )
3. Accuracy of instrument (Add known unbalance and see if machine gives correct change in total )

Secret Number 6. Payload should usually weigh at least 2% of maximum capacity of machine.

You know you shouldn’ t measure your weight on a 20,000 lb truck scale. The same principal applies to balancing a 4 pound payload on a 1000 pound spin balance machine.

MOI accuracy is reduced when the MOI of a payload is smaller than the tare MOI of the instrument. For example, if a payload has a MOI of 10 lb-in2, and the instrument has a tare MOI of 10,000 lb-in2, then a 0.1% change in tare due to an ambient temperature change will result in a 100% error in the measured MOI of the payload.

CG accuracy is generally expressed in terms of moment accuracy. If a 10,000 lb-in full-scale CG instrument has a moment accuracy of 1 lb-inch, then this will result in a CG error of 1/10,000 = 0.0001 inch for the maximum payload weight of 10,000 pounds. If the payload weighs 1000 pounds, then the same instrument will give an error of 1/1000 = 0.001 inch.  If the payload weighs 100 pounds, then the error will be 0.010 inch, and so on.

POI accuracy is generally expressed in terms of minimum detectable unbalance in lb-in2. The bigger the machine, the larger will be this minimum detectable POI. The reason for this increase is that the larger machine must use a heavy rotating spindle. Minor alterations in the rotational center of this spindle will cause a change in the residual unbalance of the machine. Furthermore, the heavy spindle has a larger change in bearing force for a given displacement magnitude of ground vibration. If the machine spindle weighs 1000 pounds and you are trying to balance a payload weighing only 1 pound, then ground vibration of only a few millionths of an inch may affect your accuracy.

Secret Number 7. You must define your symbols and polarities.

Define which axis is X, which axis is Y, etc. Your X may be someone else’ s Y. Even within one company, one department may call the roll axis X and another department may call it Y. If you submit the data without defining the axes, each group will use its own set of coordinates in interpreting the data. These problems can be minimized by adopting a standard coordinate system for mass properties measurement. See my 1991 SAWE paper entitled “Proposed SAWE Standard Coordinate System”. We are planning to discuss this issue at the 1992 SAWE conference and adopt a standard. Hopefully, by the time you read this paper, such a standard will be in place.

Moment of inertia can only be positive, so there is never any confusion regarding sign. However, you should determine whether this magnitude should be expressed through the geometric centerline of the vehicle or through its CG about an axis parallel to the geometric centerline or rotated so the data is through the principal axes. In most cases, there will not be a big difference in these three magnitudes. This can lead to confusion, since it will not be immediately obvious that the wrong data is being presented.

Center of gravity can be positive or negative. You should determine whether your positive axis agrees with the definition of axes used by the recipient of your data. Furthermore, CG distance can be expressed along a coordinate system defined by the geometry of the vehicle or along the principal axes.

Product of inertia can also be positive or negative. Since this quantity is derived by multiplying the incremental masses by two different distances, the POI sign is even more prone to error than the sign of the CG data. I frequently hear the comment: “I can calculate POI, but I never get the sign right”.  What usually happens is not that the sign  is wrong, but that the mass properties engineer and the recipient of his data are using different coordinate systems.

About vs along Moment of inertia is expressed about an axis. CG can be a distance  along an axis, or a moment about an axis (CG along X corresponds to the CG moment about Y). POI is relative to two axes (or it can be a tilt angle in a plane defined by two axes).

Testing the Mass Properties Data for Reasonableness Without knowing any of the details of a particular payload, it is possible to check the measured data to see if it falls within some basic guidelines:

CG No matter how the mass is distributed, the CG of an object must be within the outline of the object. If the object is a thin rocket whose nose is 100 inches from the aft end, then the CG certainly must be less than 100 inches from this aft end. It’ s likely that the CG will be between 33 inches and 66 inches from the aft end. The radial CG will normally be close to center (within 0. 100 inch).

MOI The MOI about one axis cannot be greater than the sum of the MOI’ s about the other two orthogonal axes. A proof of this handy test is found on page 12 of George Strom’ s SAWE paper No. 1946 entitled “Matrix Methods for Mass Properties”.

The radius of gyration of the payload must be less than the longest dimension at right angles to the axis. For example, if the MOI about the roll axis of a 100 pound missile is measured to be 10,000 lb-in2, then the corresponding radius of gyration is the square root of 10,000/100 = 10 inches. If the radius of the surface of the missile is 5 inches, then there is something wrong with the measurement. If the missile were of constant density, then you would expect the radius of gyration to be about 70% of the radius, so in this case the true radius of gyration might be about 3.5 inches.

GIVEN: Body with moments of inertia – I_xx, I_yy, Izz products of inertia – I_xy, I_xz, I_yz

COMPUTE:

If the moments and products of inertia represent a real body, the following four inequalities must be true.

A² ,c: 1

B² ,c: 1

C² ,c: 1

A² + B² + C² – 2 ABC ,c: 1

If any of the four inequalities is not true, then rotating the moments and products to the principal axes will yield one moment of inertia greater than the sum of the other two or one moment of inertia negative.

Conclusions Considerable knowledge and skill is required for you to get accurate mass properties measurements. If you buy a good instrument, then the limit on your accuracy will be one of the seven factors discussed in this paper.

About the Author

Richard Boynton holds a B.E. degree in Electrical Engineering from Yale University and has completed graduate studies in Mechanical Engineering at Yale and M.I.T. He is the author of over 51 papers, including 18 papers presented at past SAWE conferences. Mr. Boynton has been a member of SAWE for 24 years and is currently President of the Boston Chapter. He has designed many of the mass properties measuring instruments manufactured by Space Electronics. Also, Mr. Boynton is the Chief Executive Officer of Mass Properties Engineering Corporation and is a professional folksinger.

This paper by Richard Boynton was originally published for presentation at the 51st Annual Conference of the Society of Allied Weight Engineers, Inc., May 18-20, 1992.  Paper No. 2095, Index Cat. No. 6. Click here to download a pdf copy. 

The post The Seven Secrets of Accurate Mass Properties Measurement appeared first on Raptor Scientific.

The two primary factors affecting selection of a Space Electronics Gimbal Balance Machine are weight of the gimbal and its unbalance tolerance.

Maximum gimbal weight.   The machine must accommodate the maximum weight and size of the gimbals to be balanced. However, it should be no larger than necessary since large machines are more expensive and less sensitive than small machines.

Sensitivity to unbalance is defined by Space Electronics as the smallest unbalance moment that can be repeatably measured for a gimbal with a rotation angle of ±45^o. If the gimbal rotation angle is smaller than this, sensitivity of measurement will be reduced. The table below illustrates this effect.

You should select a Gimbal Balance machine which is at least 5 times more sensitive than the balance tolerance for the gimbal. To calculate the rated machine sensitivity (S) required when given a balance tolerance (G), use the following formula:

S = 1.4(G)(F)SIN(a)

where a is the maximum rotation angle of the gimbal (measured from the midpoint of rotation) and F is the ratio between machine sensitivity and gimbal balance tolerance requirement (typically 0.2).

EXAMPLE If the gimbal must be balanced to a tolerance of 10 g-cm, a maximum of 20% of the allowed unbalance may be used up by the machine sensitivity limit, and the gimbal rotation angle is ±25^o, then G = 10 g-cm, F = 0.2, a = 25^o. The required machine sensitivity:

S = 1.4 x 10 x 0.2 x sin 25^o = 1.2 g-cm

Therefore, a machine with a sensitivity better than 1.2 g-cm should be purchased (such as the Space Electronics Model GM904 with rated sensitivity of 0.5 g-cm).

Optional automated gimbal rotation.  The measurement process requires that the gimbal be rotated precisely to four positions. This can be done by hand. However, if the machine is to be used to balance a large number of gimbals, then it is convenient to have this rotation accomplished electrically. In addition to its convenience, this option generally improves the sensitivity of the measurement, since electrical rotation has less disturbing influence on the instrument than manual rotation. Space Electronics can provide the electronics necessary to do this automatically via commands from the computer. This option includes the software necessary to interface with the drive electronics provided by the customer. It also includes a beryllium copper flex strip to provide power to the gimbal without reducing the sensitivity of the measurement. We can also provide a means of electrically reading the actual rotation angle. This true angle is then used in the calculation of gimbal unbalance.

The post Choosing the Right Gimbal Balance Machine appeared first on Raptor Scientific.

Figure 1 – The Gimbal is rotated to four different positions.

• All measurements are made in one machine setup, which allows unbalance about both rotation axes to be determined simultaneously.

• Unbalance moment is reported directly in terms of gimbal coordinates.

• The gimbal is static during measurement, eliminating the effects of bearing friction and the possibility of damage which can occur using dynamic methods .

• Fully assembled gimbals (complete with wiring) are tested, significantly improving accuracy over methods which balance individual components.

• Our force rebalance technology provides highest sensitivity combined with ruggedness and high speed.

• Optional custom software can provide a printed report of optimized weights and locations required to reduce unbalance to acceptable levels. Trial and error balancing is eliminated.

Position 2

Position 3

Position 4

Custom weight correction software After the  machine measures the unbalance about the gimbal axes, it is necessary to choose the proper ballast weights required to bring this unbalance within tolerance. If the ballast weight locations were directly in line with each axis, then this would be a simple matter. However, usually the locations do not fall directly on an axis, resulting in a moment change in more than one axis when a weight is added to the gimbal. This interaction makes the balancing solution very difficult to solve. If this is done by trial and error, it can take as much as 8 hours to balance a single gimbal. Raptor Scientific can provide custom software that requires less than 10 minutes to examine all combinations of available weights and locations and presents the best solution. This software is unique to each gimbal. If you are interested in a quotation, we will need specific information about your gimbal. Contact our sales department.

To balance the gimbal, the operator mounts one 1736AX weight at location 1, two 1738A weights at location 3, etc. After mounting correction weights, the operator re-measures the gimbal to verify that it is within unbalanced tolerances.

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Absolute vs. Relative Measurements

“Should I take absolute measurements, or should I compare my blades to a master blade? What difference does it make?” Let our chief engineer, Kurt Wiener, answer:

“Both methods are used, but the accuracy will be greater with relative measurements. For small blades, the difference is not significant because of the proximity to absolute zero. But when it comes to large blades such as fan blades, we strongly recommend using relative measurements. Using the same scale, the accuracy of an absolute measurement is in fact typically 4 to 5 times worse than that of a relative measurement.

“If a set of blades is being sorted to achieve balance on an engine, then it is only important that the measurement be relative to a master blade. In other words, you are looking for the difference between blades. Since all measurements are made using the same setup, this will result in greater balance.”

“However, if blades are being sorted for later use, and they may be combined with blades measured using other fixtures or moment weighing machines, then it is more important that the true moment value be measured. It is much more difficult to make an absolute measurement of blade moment than a relative measurement. A major source of absolute measurement uncertainty is in the fixturing used. The contact between the fixture and the blade will determine the absolute moment measurement accuracy.”

Sensitivity to Temperature and Drafts

Measuring blades in an open environment is a challenge for conventional moment weight scales. Temperature variations and drafts induce large errors and make it impossible to take repeatable measurements. Space Electronics moment eight scales were specially designed to address these issues. The transducers used in Space Electronics moment weight scales are temperature compensated.

Space Electronics moment weight scales come with optional shrouds that cover the blades during measurement. Alternatively, you may want to have a temporary wall around your moment weight scale to protect it from drafts while measuring turbine blades.

Measuring at engine radius vs. optimal measurement radius

Most turbine drawings specify measuring blades at engine radius. However in some cases, in particular for most power generation turbine blades, this is not practical because the engine radii are large and the blades short. Measuring at these radii causes much more rocking inertia and sensitivity to windage than measuring at a short radius would generate.

Our moment weight scales allow measurements at any radius. The results can be translated to engine radius by the computer system. The only disadvantage to measuring at a different radius is the need to measure each blade weight (this is called pan weight) very accurately, in order to perform the conversion from measurement radius to engine radius. This means two measurements per blade, which takes more time.

Advantages of Precise Measurements

By following these easy balancing tips you ensure that your blades are measured more accurately. The blade distribution software is more likely to find a better-optimized balancing solution, with a lower residual unbalance. Your rotor will be better balanced, and therefore easier to integrate with other rotors when re-building the turbine.

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1. Why impose a limit on moment of inertia?

Moment of Inertia (MOI) is a measurement of a clubhead’s resistance to twisting. It is a strong indication of the “forgiveness” of a clubhead – that is, the extent to which a good result can be achieved from a less than ideal contact with the ball. Further increases to MOI could reduce the challenge of the game by reducing the skill required to hit the ball straight. In addition, that could also result in an increase in average driving distance by reducing the likelihood that swinging faster will produce a poor result.

As stated by the United States Golf Association in a communication to golf manufacturers in March 2005, “moment of inertia of driver heads has approximately tripled over the past 15 years. The USGA is concerned, however, that any further increases in clubhead moment of inertia may reduce the challenge of the game. Future materials with greater strength and lower weight than materials currently used in clubheads could potentially enable significant further increases in moment of inertia. There may be other means of further increasing moment of inertia as well.”

Research conducted by the USGA has shown that the clubhead size limitations already in place will not effectively prevent increases in clubhead moment of inertia beyond the levels achieved by clubs submitted to the USGA in 2005. The USGA has allowed substantial increases in MOI, but it now believes that a limit is appropriate.

2. How is the rule enforced?

Every new club needs to be submitted to the USGA to test conformance with the Rules of Golf. Every club that has been submitted to the USGA and already been ruled conforming to the Rules of Golf by the USGA remains conforming with the limit on moment of inertia.

Every new club as submitted by the club manufacturer will be tested for MOI around the vertical axis through the clubhead CG. Clubs with movable weight designs need to meet MOI limits in all intended configurations.

The USGA will use its own test rig to measure a clubhead. If the measured MOI is under 6,000 g-cm2 (5,900 + 100 for tolerance), the clubhead passes the test. If it is over, it fails the test. Please note that the USGA will only pass or fail a clubhead, without giving any justification.

3. How can I make sure that my clubhead will pass the USGA test?

Before we can answer this question, you need to understand how to measure moment of inertia through the center of gravity and how accurate the results are.

Moment of inertia is measured by making the clubhead oscillate around the axis you need to measure. MOI is proportional to the period of oscillation. In our case, the USGA needs to measure the moment of inertia around the vertical axis passing through the center of gravity of the clubhead. As the center of gravity is unknown, the USGA makes 6 measurements around 6 axes parallel to the axis they want to measure. Through a least squares regression analysis, they calculate the moment of inertia around the vertical axis passing through the center of gravity.

Any measurement has an uncertainty. The USGA uses an instrument that has 0.5% uncertainty on each measurement. However, the method used to recalculate MOI through CG gives different accuracies depending on several factors, including the size of the clubhead and the CG offset of the clubhead from the MOI instrument centerline. In reality the measured MOI can have an uncertainty of up to 92 g-cm2.

What does it mean for your clubhead? It means that if the real MOI of your clubhead is 6,000 – 92 = 5,908 g-cm2 then you can be sure that it will pass the USGA test.

4. How can Raptor Scientific help me?

Raptor Scientific has been working with golf club manufacturers for more than 40 years to measure moment of inertia and center of gravity of golf clubheads and balls.

Raptor Scientific offers a wide range of products and services to meet all needs and all budgets:

• Moment of inertia measurement instruments
• Center of gravity measurement instruments
• Measurement services (moment of inertia and center of gravity)
• Fixture design and manufacturing to reduce measurement uncertainty
• Consulting in mass properties
• Automated least squares regression analysis as defined by the USGA to calculate the moment of inertia through the center of gravity (contact us for more information)

5. How to optimize a golf clubhead design to maximize moment of inertia while staying within the USGA limitation

So how can we make sure that a particular clubhead complies with this requirement? It needs to be measured. And here comes into play your own ability to measure moment of inertia. If you measure MOI with a 0.5% accuracy instrument, you have to leave the same margin for your own measurement. It means that you can only design clubheads with a moment of inertia of 5,908 – 92 = 5,816 g-cm².

Raptor Scientific can help you achieve better designs. Our different models of moment of inertia instruments can measure moment of inertia with an uncertainty of 0.01% to 0.25%. Here are the results achieved with these instruments.

At this point we believe that the moment of inertia of clubheads is far enough from the limitation imposed by the USGA that an XR10 gives results that are accurate enough. However, getting closer to the limit will become a concern with the advances in technology.

The post How to Comply With the USGA Ruling on moment of Inertia of Golf Clubheads and Optimize Clubhead Design appeared first on Raptor Scientific.

Satellite flight can be divided into four phases:

• Initial launch, using a two or three state booster
• Injection into the approximate desired orbit, using a “kick” motor on the satellite
• Fine positioning to set the exact initial orbit or to correct for atmospheric drag, using small thrusters on the satellite
• Attitude control to rotate the satellite about its CG so it points in the designed direction (many methods are discussed on this page)

Initial Launch

Mass properties play a key role in the first few seconds of rocket flight. At the initial launch, there is no forward motion of the rocket and consequently no aerodynamic forces to stabilize the rocket motion. Alignment between the center of thrust and the center of gravity of the vehicle is critical to prevent large offset moments from being generated. If an engineer is used to dealing with flight vehicles that are inherently stable (such as aircraft), then the sensitivity of rocket flight to CG misalignment may come as a surprise. It is not uncommon for engineers to specify tight tolerance for rocket CG, such as +/- 0.002 inch. Unfortunately, many companies measure the rocket CG in a dry condition (i.e. without fuel). Since the weight of fuel can be as high as 85% of total vehicle weight, making a dry measurement can be almost meaningless unless you subsequently measure the same rocket after fueling. In my opinion, the only reliable way to measure rocket CG is to locate the measuring equipment in the fueling area, and make the measurement (and perform ballasting if necessary) after the rocket has been filled with fuel.

For aircraft ( and air-to-air missiles), the center of pressure of the airframe is aft of the center of gravity of the vehicle, so that aerodynamic forces tend to align the vehicle to the direction of flight. This is not the case for the first few seconds when a rocket is launched from earth, since there is no initial forward velocity to create aerodynamic forces. The aerodynamic alignment disappears again when the rocket leaves the denser atmosphere. If the rocket CG is offset from the center of thrust, then a torque will be created equal to the CG displacement D times M*(g+a). The gimbaled rocket motor will rotate to align the thrust with the CG, causing the rocket to tilt relative the direction of flight. As the rocket builds up speed, this causes a drag. Subsequently tilting the rocket motor to align the rocket aerodynamically will reduce this drag but will waste a portion of the thrust force to correct for CG offset.

As fuel is consumed, the rocket CG changes dramatically. To minimize flight problems, the fuel tanks are nominally located exactly on the centerline of the vehicle, and the tanks contain pressurized bladders or baffles that minimize sloshing of the fuel and attempt to keep the mass of the fuel centered. However, if the fuel tank is not centered properly relative to thrust, as the fuel is consumed, the CG will shift along the yaw or pitch axis due to the increased importance of the mass of structure. The solution to this problem is to measure the rocket CG without fuel and then add the fuel. If the tanks are centered properly, the CG should not shift appreciably in a radial direction when fuel is added. CG will most likely move along the roll axis, but this is of a smaller consequence since the rocket does not use aerodynamic forces for stability, so that the distance along the roll axis between the CG and the aerodynamic center of pressure is not a major factor.

It is easy to measure CG location relative to hard points on the rocket structure, using commercial equipment such as the instruments manufactured by Raptor Scientific, but often it is hard to accurately determine the relationship between CG and the center of thrust. What reference do you use for the thrust centerline? Is the center of thrust aligned with the dimensions of the rocket motor exit cone or is it canted slightly due to any number of complex factors within the engine? What about tolerance buildup between the mounting surface of the rocket on the center of gravity instrument and the location of the rocket motor nozzles? Thrust alignment is determined by mounting a similar rocket motor in a “thrust stand” that is instrumented with strain gage load cells and determining thrust alignment (as well as total thrust). Optical means can be used to determine alignment of the motor with the rocket structure. Photogrammetric techniques have been refined so that three axis positioning can be measured with a few thousandths of an inch.

In the early days of satellite launches, it was not uncommon for a rocket to make a sharp turn shortly after leaving the launching pad and have to be destroyed. Other rockets would begin to “pinwheel” (tumble end over end) shortly after launch. I was hired as a consultant by a government agency in a foreign country after a rocket turned around and flew into the launching pad, creating a 6-foot deep hole in the concrete and destroying the launch facility. One of the major causes of these failed flights was CG misalignment with the center of thrust. Another was the degree of balance of the gimbaled guidance system (see SAWE Paper No. 3320 Static Balancing a Device with Two or More Degrees of Freedom –(The Key to Obtaining High Performance On Gimbaled Missile Seekers). Now there appears to be a much greater awareness of the importance of aligning rocket CG and of some of the pitfalls in the process.

Rocket Moment of Inertia

Although moment of inertia is less critical than center of gravity, it does have a significant effect on flight. At the instant of lift off, transverse (i.e. pitch or yaw) MOI is the only “force” resisting the tilting of the rocket. This is not a new concept—ancient tribesmen discovered thousands of years ago that a spear flies straighter if it is long and narrow. More recently engineers have assumed that they could calculate moment of inertia with sufficient accuracy that they didn’t need to measure it. I have seen reports that listed the expected MOI error as “less than 2%”. In fact these calculated MOI values were in error by more than 100%. The reason is that fluid makes up about 85% of the rocket mass, and the assumptions the mass properties engineers were using were erroneous. One school of thought assumed that the MOI of the fluid in the tanks was zero, since the fluid would remain stationery and the tank would move about it. The other school of thought assumed that the fluid should be treated as a solid. Neither of these assumptions is correct. As I have demonstrated in two SAWE papers on fluid dynamics, the effective MOI of the fluid depends on the shape of the tank and the presence or lack of baffles within the tank. {See SAWE PAPER No 2459 The Moment of Inertia of Fluids, and SAWE Paper No. 3006 The Moment of Inertia of Fluids—Part 2.}

Some of the conclusions of these papers are summarized below:

1. Roll Moment of Inertia The MOI of the fluid in a cylindrical tank is not zero about its centerline as is commonly supposed. For a simple right cylindrical tank with flat ends which are perpendicular to the centerline, it ranged from about 3% of the solid equivalent MOI for a fluid with the viscosity of water (or hydrazine) to 93% of the solid equivalent MOI for a fluid with the viscosity of gear oil. Since the mass of the fluid can be over 80% of the mass of the vehicle, this can be a significant contributor to the total MOI. In fact, the liquid can be the largest contributor to roll moment of inertia (as well as pitch and yaw MOI).
2. Pitch and Yaw Moment of Inertia The MOI of the fluid in a cylindrical tank about an axis perpendicular to its centerline typically ranges from 40% to 80% of the MOI which would result if the fluid were solid. The percentage of mass which should be used to calculate this effect depends on the aspect ratio.
3.  Viscosity As the viscosity of the fluid increases, the fluid moment of inertia increases. This effect is most dramatic for the roll MOI about the centerline.
4. Other Effects of Fluid If a task spins continuously in one direction, eventually all the fluid within the tank spins with the tank. For a given deceleration force acting on the spinning tank, an empty tank will slow down more quickly than a filled tank. In our tests, if the tank was not completely filled, its speed decreased and increased in an oscillatory fashion as it gradually slowed down. I would expect different results at zero g.

Rocket Product of Inertia

Consider the rocket shown along the left margin of this page. The red weights have been added to simulate a product of inertia unbalance. Since they are equal in mass and located equidistant from the CG, the CG position remains unchanged. However, when the rocket tilts about its CG to change the direction of flight, a couple is created, causing the rocket to rotate. The rocket will want to rotate about its principal axis, resulting in a coning of the rocket and increased drag.

Spin stabilized rockets will tilt to align with the minor axis along the length of the rocket. This results in an “angle of inclination”. Given a certain product of inertia, the amount of tilt (“A” degrees) is related to the moment of inertia difference between the major and minor axes by the following formula, where Pxy, Iyy, and Ixx are in the same units.

Second Stage

Most satellite launches require at least two stages. When the first stage runs out of fuel, explosive bolts release the second stage, which then ignites. Second stage motors may use a different technology than the first stage. Because of the greatly reduced mass, this second stage should have a lower thrust to prevent excessive g load on the satellite. Since the second stage is not flying in a dense atmosphere, it does not have to have an aerodynamic shape, and there is little concern that protruding objects will be damaged by windage forces.

The second stage will have enough fuel to put the satellite into orbit, but this may be a temporary orbit that has to be modified by the rocket motors in the satellite.

Injection into the Approximate Desired Orbit

Achieving orbit depends on two factors:

1. Reaching the desired altitude above the earth
2. Achieving the minimum velocity parallel to the earth to sustain orbit

The speed needed to orbit the earth depends on the altitude, according to Kepler’s formula:

V = sqrt ( g * R^2 / (R + h) )

where V is the velocity for a circular orbit, g is the surface gravitational constant of the Earth (32.2 ft/sec^2), R is the mean Earth radius (3963 miles), and h is the height of the orbit in miles.

The higher the altitude, the lower the required velocity parallel to the earth. For example, to orbit the earth at an altitude of 100 miles, a velocity of 17,478 MPH is required. A geostationary orbit at an altitude of 22,236 miles requires a velocity of only 6877 MPH. At this height above the earth, the satellite orbits the earth once in 23hrs 56 minutes 4 seconds–the same rotational period as the earth, so the satellite constantly remains overhead of a fixed point on earth. (For you non-astronomers, the rotational period of the earth is not 24 hours. The extra 3 minutes and 56 seconds on an earth-based clock compensates for the rotation of the earth about the sun, so that the sun is always directly overhead at noon.)

Establishing the Final Exact Orbit

Generally, the result of launch is that the orbit is elliptical rather than the desired circular orbit. An additional step is then required to “kick” the craft into a circular orbit. By firing a rocket motor when the orbit is at the apogee of its orbit (its most distant point from Earth), and applying thrust in the direction of the flight path, the perigee (lowest point from Earth) moves further out. The result is a more circular orbit. Small “vernier” rocket motors called thrusters can then be used to precisely position the satellite. These attitude control rocket motors for satellites and space probes are often very small, an inch or so in diameter, and mounted in clusters that point in several directions.

Importance of Aligning Thruster Center of Thrust with Satellite CG

If a thruster is to be used to translate the position of a satellite, then it is essential that it act directly through the CG of the satellite. Otherwise the satellite will spin rather than move in a straight line.   It is often difficult to establish the relationship between the nozzle centerline and the CG measurement of the satellite. A novel method is described below, using a special test weight that is inserted directly into the nozzle cone.

The “Boynton Method” used to measure CG to thrust centerline directly is:

• Determine the center of thrust of the steering rocket motor. This may not be exactly in the center of the nozzle, particularly if the rocket motor body is at right angles to the nozzle. This quantity can be determined by mounting a motor on a rocket thrust stand such as those manufactured by Raptor Scientific. The rocket motor is then fired and the thrust misalignment is measured.
• Construct a precision calibration weight which fits precisely in the nozzle of the thruster and whose CG is at the nominal center of thrust of this type of motor.
• Measure the CG of the satellite with this weight in place. Remove this weight and make a second measurement. Then remove the satellite from its test fixture and measure the tare CG. By subtracting the tare measurement from the measurements of the satellite, the CG position of the satellite can be determined. By subtracting the measurement of the satellite from the measurement with both satellite and calibration weight, the CG position of the precision weight can be determined. If the CG of the satellite is in the same location as the CG of the weight, then the thrust is aligned with satellite CG. If they are different, then this difference is the thrust misalignment error. This method takes into account all the many dimensional errors that can exist in the mounting structure of the thruster. It measures what you want to know directly.

In order for this method to work, it is necessary for the nozzle weight to be much larger than the opening in the nozzle. We have made these weights with a tooling ball which fits in the nozzle, attached to extension arms to support the weights. The total structure is adjusted so its CG is precisely at the center of the ball. In this way, the weight can be placed in any orientation relative to the nozzle. {This process is described in more detail in SAWE Paper No. 2174 “Measuring the Mass Properties of the Brilliant Pebbles Satellite” by Richard Boynton.}

Fuel in Tanks and Pipes

In this illustration, the fuel tanks and the piping are significantly offset from the centerline of the vehicle. To evaluate the effect of fuel mass, it is recommended that the satellite CG and MOI (and in some cases POI) be measured with the satellite dry and then again with both the tanks and the fuel lines filled with fuel. This will provide data for the change in mass properties as fuel is consumed in flight. If the design is ideal, the CG will not shift significantly when fuel is used up (the MOI will always decrease). Fuel mass is a major factor affecting mass properties for many satellites and cannot be ignored.

Attitude Control

The previous discussion has been concerned with position control — placing the satellite into the desired nominal orbit. Attitude control concerns the angular orientation of the satellite.

For orbiting satellites, generally the position accuracy required is not particularly high, and in fact the satellite will gradually drift from its position due to the small drag from the very thin atmosphere, solar wind against the solar panels, etc. Weekly or monthly boosts may be required from thrusters on the satellite to restore its position. The Hubble Telescope drops back toward the earth by about 150 feet a day, but has no booster, because the resulting gas would surround the telescope and might coat the mirror with a thin film, blurring the image. If the Space Shuttle doesn’t push it back into orbit when it makes its planned repair trip, then the Hubble Telescope will burn up in the atmosphere sometime between 2010 and 2030.

In contrast with position control, attitude control requirements can be extraordinarily high. For example, the Hubble Telescope requires a pointing accuracy of 0.007 arc second (2 millionths of a degree)—equivalent to the width of a human hair at a distance of one mile, or getting a hole in one when driving a golf ball from the east coast of the USA to the west coast.

Forces disturbing a satellite’s attitude Even though a satellite in orbit is flying an almost pure vacuum, there are some subtle forces acting on it that disturb its position over time. Any magnetic objects in the satellite are attracted to the earth’s magnetic field, which varies as the satellite orbits the earth. Gravitation attraction also varies, since the earth is not a perfectly round sphere. These effects are particularly important for lower altitude orbits (i.e. 200-400 miles) where many spy satellites are located. The Hubble Telescope is also located at a low orbit so it can be serviced by the Space Shuttle. Another disturbance in low earth orbit is the residual atmosphere dragging on the solar panels.

At a high orbit, such as the geostationary orbit at 22,236 miles, these effects are smaller. However, solar wind acting on the solar panels and mass position changes within the satellite itself due to rotation of antennas or telescopes can disturb the attitude of the spacecraft.

Reasons for needing attitude control    Most satellites point either a telescope or an antenna toward earth. As the satellite makes its way around the earth, it must rotate to continue to face the earth. Furthermore, if the satellite uses solar power, it must turn its solar panels so they continuously face the sun. Unless some method is employed to control attitude, the rotation of the mass of the solar panels will produce an opposing rotation of the satellite.

The following sections describe different methods of attitude control. Many satellites use more than one of these methods.

Passive Attitude Control Most current satellites use closed-loop servomechanisms to maintain pointing angle. However, in the early days of the space program, two open-loop methods were used to control attitude: spin stabilization and gravity gradient stabilization. There is a renewed interest in some of these methods.

Gravity Gradient Stabilization is the simplest of all methods. Basically it consists of providing a long thin structure that is large enough so that the force of gravity is significantly larger at the end closest to the earth, so that the spacecraft then spontaneously orients itself so that the axis of minimum MOI points toward the gravitational center of the earth. No fuel or apparatus is required. This stabilization technique works well for the earth’s moon, since it is large and rigid. It has not worked well for artificial satellites, because there is essentially no damping in space and the satellite oscillates like a pendulum (the greater the ratio of minor to major MOI, the longer the period). One solution has been to attach a thin cable to the satellite and tether a mass toward the earth. A closed loop control mechanism can then vary the effective attachment point and damp the oscillations (but the major advantage of being passive is lost). The Chinese have been experimenting with adding an aerodynamic stabilizer for satellites in low earth orbit, where gravity gradient is higher and atmosphere is denser. Another method to overcome the oscillations is to add a viscous damper, a small can or tank of fluid mounted in the spacecraft, possibly with internal baffles to increase internal friction. Friction within the damper will gradually convert oscillation energy into heat dissipated within the viscous damper. As this system has two stable states, some way is required to flip the satellite and its tether end-for-end if needed.

Spin Stabilization is an ancient technology—hundreds of years ago gun manufacturers discovered that spiral grooves in a rifle barrel would cause the bullet to spin and improved the accuracy. The first American satellite—Explorer1 in 1958—spun around its long axis. In addition to defining its position, this spin was supposed to create centrifugal force to make the wire antennas extend out from the body of the spacecraft. Much to the surprise of JPL engineers, the spun object began to cone after launch and was soon tumbling end over end. What the engineers hadn’t realized was that bullets obtain stabilization because their center of pressure is aft of their CG. An object in space has no aerodynamic stabilization and will rotate about either its minor or major axis, but it favors the axis of major MOI because this axis results in minimum kinetic energy. We have a demonstration of this at our Mass Properties Seminars at Raptor Scientific. We rotate a yoke in a spin balance machine with a long cylinder pivoted at its CG. The cylinder is spun in a vertical orientation about its long axis. It maintains this position for a short time and then suddenly flops over into a horizontal orientation.

Spin stabilization is still used for certain types of spacecraft. The rules for this technique are:

1. Spin about the major principal axis, not the minor axis. This means that the satellite must be short and fat (not convenient for launch, since the rocket is long and thin).

2. Dynamically balance the spacecraft so that the product of inertia is small. The angle of inclination of the spacecraft is a function of the difference between the major and minor MOI. If you want to stabilize the vehicle and have it resistant to the effects of product unbalance, make the difference in moment of inertia as large as possible.

A variation on the spin stabilization technique is to use a two section satellite, connected by a bearing. One section spins to provide stability. The other is stationary and adds to the axis MOI to make the spun axis into the principal axis with highest MOI. One feature of this design is that is fits conveniently into the outline of the launching rocket.

Active (Closed-loop) Attitude Control

The techniques discussed previously provide enough attitude control to allow a broad band antenna to point toward the earth. However, they are not accurate enough for applications involving a telescope or a parabolic reflector. Closed loop control requires three elements: a sensor to determine the current attitude, a computer to determine the error between the desired and actual attitude, and a torque device commanded by the computer to change the attitude of the satellite.

Since this is a three dimensional problem, usually at least three sensors are required. These can be gyroscopes, a sun sensor, an earth horizon sensor, a moon sensor, a star tracker, etc.

Obviously, sensors for astronomical instruments such as the Hubble Telescope have extraordinary accuracy, whereas as spy satellites that scan large areas without specific pointing can use less accurate devices.

The computer must be programmed to command the torque devices to correct the pointing error. Like all servomechanisms, overshoot and instability can be a problem. In order to anticipate any problems in orbit, Raptor Scientific has developed “space simulators” that float the satellite on a thin film of air, simulating the frictionless environment of space. These are discussed in SAWE Paper No. 2297. Using a Spherical Air Bearing to Simulate Weightlessness by Richard Boynton.

The torque devices can be small rocket motors (“thrusters”), magnetic devices that couple to the earth’s magnetic field, or momentum wheels that make use of Newton’s laws of conservation of momentum. These devices are discussed in the following sections.

Attitude control using small thrusters In the previous discussion of satellite position, I emphasized the importance of thruster alignment with the satellite CG. For attitude control, it is important that a thruster not be aligned with the satellite CG in order to create a turning moment. Micro-thrusters can be used to intentionally create a moment in order to realign the satellite or to unload a momentum wheel (see discussion of momentum wheels in later section). Typically an attitude control thruster supplies a few millisecond pulse of energy. The effect is like tapping the satellite with a small plastic hammer. A few seconds later, an opposing tap is given to stop the satellite in its new position. The simplest thruster type uses compressed gas such as nitrogen. This is not very efficient, since the specific impulse (exhaust velocity) is low compared to other methods, so that a relatively larger volume and mass is required. A more common type of attitude control thruster uses hydrazine as a fuel.  The hydrazine is controlled by a solenoid valve, which emits a millisecond pulse of hydrazine to a nozzle containing a catalyst. The hydrazine spontaneously ignites when it comes in contact with the catalyst.

Thruster control is essentially a digital process. The smaller the pulse, the finer the control can be. One serious drawback of this method is that it requires fuel to accomplish its goal. When the fuel runs out, the satellite has reached the end of its lifetime. This is not true of reaction wheels, which get their power from the sun (discussion of this technique follows).

Another drawback of thruster control is that the mass properties of the vehicle change as the fuel is consumed. If opposing thrusters are fired in pairs, CG shift can be minimized, and pure rotation results. If hydrazine is used as a fuel, the MOI of the vehicle is dependent on temperature, since the fuel becomes more viscous at low temperatures.

Magnetic Stabilization The earth’s magnetic field can be used to steer a satellite. Long steel rods wound with fine wire form electromagnets that are used to pull the satellite toward the earth. The polarity can be reversed to push the satellite away. Since these are offset from the satellite CG, they cause the satellite to rotate.

Momentum Wheels This is one of the most elegant and accurate methods of attitude control. Momentum wheels take advantage of the laws of conservation of momentum. Changing the speed of a wheel causes the satellite to respond by turning in the opposite direction to the change. Since the momentum wheel has a very small MOI relative to the satellite, and its speed can be controlled with digital precision, extremely fine attitude control is possible.

In its best implementation, four wheels are used. Three are mounted at right angles, so that each axis of the spacecraft can be controlled independently. The forth is oriented so that it can be used to replace any of the other three in the event of failure. In this instance, there will be an interaction between axes, requiring more complex control.

If the speed of the wheel is increased in a particular direction, the satellite will rotate in the opposite direction. Conversely, decreasing the speed of the momentum wheel causes the satellite to rotate in the same direction as the momentum wheel. Absolute speed does not have an effect on satellite position; it is only the change in speed that matters. Theoretically, the momentum wheel can be at a standstill and then be made to rotate in either direction. However, practical problems such as static friction and hysteresis favor operating the momentum wheel at a constant speed and then increasing or decreasing this speed but maintaining the same direction of rotation.

The angular momentum of an object is the product of its moment of inertia and its angular velocity:

L = Iω

In the general case, these variables are vector quantities. However, since the reaction wheel control concept incorporates three reaction wheels which are mounted at right angles to each other to coincide with the three axes of roll, pitch, and yaw, we can consider each rotational motion separately as a linear variable. There are two angular momentums for each axis: the rotation of the spacecraft about a particular axis (such as roll), and the rotation of the corresponding momentum wheel aligned with that axis. The momentum wheel is driven by an electric motor, usually powered by solar panels. Its speed is controlled by a computer. The law of conservation of angular momentum is the rotational equivalent of Newton’s second law.

Applied to a satellite in space it states that there is no net momentum gain or loss unless acted on by an outside force. The satellite has a momentum due to its moment of inertia and speed, and the attached momentum wheel has a separate momentum due to its much smaller moment of inertia and much higher speed.  Therefore,

Is *ωs1 + I_m *ωm1  = Is *ωs2 + I_m *ωm2

where Is is the MOI of the total spacecraft (including the mass of the momentum wheel acting through the center of rotation), I_m is the MOI of the rotating disc of the momentum wheel system, ω_s1 is the initial angular velocity of the satellite before correction, ω_m1 is the initial angular velocity of the momentum wheel before correction, and ω_s2 and ω_m2 are the corresponding angular velocities after correction.

If the satellite is rotating due to disturbing influences and you want to stop its rotation, then the momentum wheel can be used to absorb the momentum of the spacecraft platform to prevent it from rotating.   In this case, ωs2 = 0.  The equation then becomes

Is *ωs1 + Im *ωm1  = Im *ωm2

or

Is *ωs1   = I_m  {ωm2  – ωm1}

Note that momentum wheels only cause rotation, not translation, and that this rotation is about the CG of the satellite.

This method works smoothly when corrections are alternately in opposite directions. However, when corrections must be made in the same direction, the wheel eventually reaches its maximum safe speed or approaches zero, at which point a “momentum dump” is required. This consists of slowing or speeding up the wheel and simultaneously employing some other means (such as thrusters or magnetic actuators) to counteract the effect of changing the speed of the momentum wheel.

Because of the extremely high speed, these wheels must be balanced to a high degree of precision to minimize vibration, which would blur a telescope’s image. These wheels are among the most likely devices in a satellite to fail, since they operate at such a high speed. Some of them use magnetic bearings, which are non-contacting. I’ve read that others use ball bearings whose balls are coated with diamond dust to minimize wear. Currently the Hubble telescope has several failed momentum wheels. Another satellite lost a key momentum wheel, but engineers were able to restore operation by revising the software to use magnetic actuators in combination with the remaining momentum wheels.

The rotational equivalent of Newton’s Second Law is that the net torque acting on an object is equal to its moment of inertia times its angular acceleration.

      t = I a

where
t is the net torque.
I is the moment of inertia.
a is the angular acceleration.

Constant angular acceleration is equal to the change in angular velocity divided by the time it takes to change.

     a = D w / D t

a is the angular acceleration
D w is the change in angular velocity
D t is the elapsed time.

Control Moment Gyros (“CMG’s”)

If a high speed flywheel is mounted in a gimbaled assembly, then the angular position of the flywheel bearing can be altered, causing the satellite to turn relative to the flywheel, whose position is fixed in space. This direct method has the advantage that the flywheel always turns at a constant speed, so that no “momentum dump” is required. The power efficiency is higher than momentum wheels. The disadvantage of this method is that precise motion of the gimbaled structure is hard to achieve. Backlash in the gimbal position drive gears, small aberrations in the gimbal bearings, and other mechanical limitation cause the gimbal position to jump unpredictably rather than move in a smooth manner. Furthermore, the mechanical structure is relatively heavy and is less reliable than other methods of attitude control. The Skylab, MIR, and International Space Station all use this concept for attitude control.

Conclusions

This paper gives a general overview of some of the methods of controlling rocket and satellite position and attitude. Since the spacecraft is unrestrained in space, mass properties define its position and motion. Evaluation of mass properties is complicated by the presence of liquid fuel, whose effective MOI is related to the shape of the fuel tank, the presence or lack of baffles, the viscosity of the fuel, and the amount of fuel remaining. Fuel is carried to the thruster nozzles via piping that often is located near the circumference of the satellite, making the effect of fuel mass in the pipes more significant in affecting both MOI and CG. The dry weight of this piping is considerably less than when filled with fuel. There’s no guarantee that the piping is symmetrical about the CG, so CG may shift when the pipes fill with fuel. Engineers are cautioned to measure rockets and satellites in a dry and wet condition, since fuel mass can be as high as 85% of total vehicle mass and can be the dominant contributor to MOI and CG location. Although mass property measurement adds another step to the fabrication process, the consequences of not making a measurement can be catastrophic—the spacecraft mission can fail, with the spacecraft endlessly tumbling end over end or not reaching its desired orbit, resulting in hundreds of millions of dollar lost, and the mission delayed for years.

If you would like the pdf version of this white paper, please complete this form to access the downloadable version.

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This is due to a phenomenon known as cross-coupling (or cross-talk). Cross-coupling is the effect that one axis has on another axis. More specifically, when one axis of a gimbal purposedly slews at a fast angular speed, the other axis reacts and rotates as well. The same phenomenon occurs when one axis reacts to vibration and the other axis reacts to it in turn. Why does cross coupling happen?

Cross coupling happens because each of the subassemblies of the gimbal has a product of inertia. To illustrate this, let us consider the following representation of a gimbaled camera:

		This camera has an unbalance.It is statically unbalanced because the center of gravity of this assembly is not at the center of the cylinder, and therefore not at the intersection of the axes. It also has a dynamic unbalance about the azimuth axis. Spinning this object about the azimuth will result in forces that tend to make it spin about the elevation axis until it finds a stable position.
		In order to compensate for the unbalance we add a weight on the opposite side. Our camera is now statically balanced (the center of gravity is at the intersection of the axes at the center of the cylinder). It is capable of looking at a slow moving target and be jitter free. This configuration eliminates most of the jitter. This static balancing process is done using a gimbal balance machine. The camera is still dynamically unbalanced about the azimuth. This object still rotates about its elevation axis when it is spun about the azimuth.
		Let us now compensate for the dynamic unbalance. We add two weights opposite of the ones that created the dynamic unbalance. Several methods can be used for dynamic balancing, including the use of spin balance machines or moment of inertia instruments to determine the cross products (also called products of inertia). The technique used depends on each gimbal. Considerations such as travel angles on each axis, mass and inertia distributions, gimbal type, rotational speed, and tolerable residual unbalance determine which method can be used. The gimbaled camera is now dynamically balanced about both azimuth and elevation. It is both statically and dynamically stable. We have eliminated the cross coupling (or cross product or product of inertia). In an ideal world the job is done.
		Unfortunately (from a balancing standpoint) most gimbaled sensors have internal adjustments. A gimbaled camera can have a zoom that changes its configuration. Or the gimbal has more than 2 axes. These provoke changes of configuration that affect balancing. The figure on the left represents a rotation of the inner axis (elevation) to track a target. Because of this rotation the outer axis (azimuth) is now experiencing a product. The camera is dynamically unbalanced about the azimuth axis again. In real situations, dynamically balancing a gimbal means finding a compromise that will lead to the best performance in a range of specified configurations.

How do I find the best balancing solution for my gimbal?

Space Electronics developed the Gimbalance by Space Electronics suite of solutions to address balancing concerns on gimbaled assemblies. It includes static balancing methods (utilizing a gimbal balance machine) and dynamic balancing solutions tailored to each gimbaled platform.

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Raptor Scientific manufactures moment weight instruments to measure and balance high-speed turbine blade for applications like jet engines, steam turbines, petroleum drilling, and pumping stations, These instruments are often called “moment weight scales“. Our instruments use a new technology which is as much as 40 times more accurate than the conventional knife-edge and load-cell technology that has been employed for the last 30 years. As a result, the moment measurement error of our instruments can be considered insignificant. This has led us to more clearly identify other sources of measurement error which appear to be widespread throughout the industry.

The problems show up in two ways: (1) a blade is replaced in the field with one of supposedly identical moment, and the engine is then found to be unbalanced; (2) a set of blades is measured at Plant A and then sent to Plant B for installation in the engine. If the blades are measured at Plant B before they are installed, the data differs from the original set of measurements. However, it often isn’ t just a simple change in scale factor (i.e the blades aren’t just 0.5% higher in moment at Plant B). There are several factors involved, resulting in what appears to be random differences. I believe I have identified the sources of these errors. This paper identifies each type of error, and gives recommendations for their elimination.

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Flutter is of great concern to any pilot, since excessive flutter has caused a number of aircraft to lose control and crash. Although any surface on an aircraft which is exposed to air flow can experience aerodynamic flutter, the most common type of flutter involves the control surfaces such as ailerons, elevators, and rudders. The mass properties of these control surfaces are critical and have to be measured with great care to make certain that flutter is minimized. Many mass properties engineers ignore product of inertia when measuring control surfaces. We suspect that these engineers will be surprised to discover that the product of inertia unbalance of the control surface can be the key element in understanding and eliminating aerodynamic flutter, and that it is vital to measure this quantity.

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Control Surface Center of Gravity

If the center of gravity of a control surface assembly is not located exactly on the hinge line, then a torque will be applied to the control surface whenever the wing (or other mounting surface) accelerates in a vertical direction. If the inertial forces acting through the CG of a control surface amplify the vibration of a wing, then the flutter amplitude will rapidly increase and can result in destruction of the aircraft. Small changes in control surface CG location can have a dramatic effect on flutter instability.

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Control Surface Moment of Inertia

The moment of inertia of the control surface is also a critical parameter. If the wing (or other mounting surface) accelerates in a rotational sense due to twist, the position of the control surface will lag behind the wing due to its moment of inertia. In other words, the orientation of the control surface will change relative to the orientation of the wing during wing twist, even if the control surface is statically balanced about the hinge line. Generally, it is necessary for control surface MOI to fall within a narrow range of values, in order to avoid flutter problems. It is not sufficient to simply minimize control surface MOI, since it is important to avoid mechanical resonant modes whose frequencies are harmonics of each other. Note: It is possible to neutralize this MOI effect by moving the CG ahead of the hingeline. There is also another consideration involving moment of inertia: the frequency of the notch filter in the actuator circuitry which drives the control surface is based on the particular moment of inertia being driven. If this moment of inertia deviates significantly from the nominal value, the control circuit can become unstable.

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Control Surface Product of Inertia

When a wing suddenly bends upward as the result of a gust of wind, mass on the outer end of the wing will experience a greater acceleration than mass near the fuselage. An aileron can be balanced statically but have a concentration of mass near the trailing edge on the outboard end and a corresponding concentration of mass near the leading edge on the inboard end. This will create an unstable condition when dynamic forces are applied.

POI can also aggravate twist of the airfoil. Although the control surface does not rotate, its rapid vibration creates inertial forces that can distort the shape of the wing.

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Summary

Performance requirements for aircraft and other aerodynamically controlled products have become ever more demanding with less and less margin for error. Materials, especially composites, have become more complex and are inherently less homogeneous than the materials previously used. This has led to a situation where the mass properties of control surfaces may not always be within the test tolerances required for safe and stable aircraft control over the full range of anticipated operating conditions. There is a growing trend among experienced aerospace and mass properties engineers to require the measurement of the mass properties of aircraft control surfaces to ensure that they meet prescribed tolerances in a flutter analysis.

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Click to download: Using a Spherical Air Bearing to Simulate Weightlessness (download now)

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The primary emphasis is to quantify, through physical principals, and verify, through experimental demonstration, the degree of static balance required to minimize the detrimental effects of external vibration to an acceptable level. The effects of dynamic balancing will also be discussed.

The principles developed for visible light optical systems carried by an Unmanned Air Vehicle (UAV) will be expanded, in a general way, to describe how these principles apply to infra red, ultra violet, and radar systems as well as variations to the requirements as a function of the vehicle on which they are mounted. This discussion will include manned aircraft, missiles, land
vehicles and watercraft.

Click here to download: Improve your Sensor Image with Balance

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The terms used in turbine blade balancing can be confusing. The two main terms used are pan weighing and moment weighing.

What Is Pan Weighing?

Pan weighing means measuring the actual mass of the blades. This would simply be called “weighing the blade” in layman’s terms. We call it pan weighing to avoid confusion with moment weighing.

Pan weighing the blades gives a first idea of how to roughly balance the rotor. By placing blades of equal mass opposite each other on the rotor, or by optimizing the distribution of mass around the rotor to have the overall average mass of the blades centered on the axis of rotation, the overall rotor should be balanced, right?

Wrong. This is only true if the average center of mass of all the blades is located at the center rotational axis of the shaft. To balance a rotor, one needs to locate the average mass of all components on the designed axis of rotation by changing blade positions, adding counterweights and/or removing material from sacrificial locations.

The moment weight of a single blade is equal to:

Radial Moment Weight = (Mass of the blade) x (Distance from the assembly axis of rotation to the center of mass of the blade)

Moment balancing of large blades, such as jet engine fan blades, involves weighing three moments of the blade — the radial, axial and tangential — which can be accomplished with the proper machine.

Real-World Examples of Moment

One of the most practical ways to describe moment is imagining a see-saw on a playground. Putting equal weights on both sides of the see-saw creates a balanced moment. Applying more or less weight to one side makes an unbalanced moment. To rebalance, you can either shift the weight positions or add or remove weight on one side.

Trying to loosen a tight bolt with your fingers is often impossible. However, a wrench makes turning much easier because the distance between the force and the pivot grows, increasing the turning moment. The further away from the pivot you apply a force, the easier the task becomes.

Static balancing — also called moment weighing — is one of the most critical initial steps for balancing a turbine rotor. High-accuracy static balancing begins with measuring the moment that each turbine blade will add to the assembly. The resulting measurements enable you to complete preliminary static balancing of your rotor quickly and precisely, showing good results for lower-precision applications and reducing the need for corrections if you are proceeding to dynamic balancing.

What Is Moment Weighing?

Moment weighing is a standard operation for turbines, being viable for power generation turbines, aircraft engines, and turbo pumps.

The following steps are necessary for balancing a rotor stage:

1. Moment weighing all the blades.
2. Determining the unbalance of the shaft and other connected hardware and entering it into the optimization software of the moment weighing scale.
3. Running the blade optimization software to determine the optimal arrangement of blades that will minimize the total unbalance of the rotor. Note that the blade optimization software is generally integrated into the computer system that controls the moment weighing scale.

Typical Applications

Moment weight scales from Raptor Scientific are ideal for both original equipment manufacturers (OEM) and maintenance and repair operations (MRO) providers in various turbo machinery applications, including:

• Jet engines: Turbine and fan blade manufacturing and turbine assembly.
• Helicopters: Blade balancing for assembly and manufacturing.
• Industrial fans: Fan blade balancing for assembly and manufacturing.
• Power generation: Steam, gas and wind turbine blade assembly and manufacturing.

We offer an extensive range of measuring solutions, from small single-axis devices for compressor blades down to 100 grams and 2 inches long to multi-axis moment weight scales for larger blades up to 200 pounds and 20 feet in length.

Discover More With Raptor Scientific

Raptor Scientific is a leading global provider of testing, measurement and engineering services, including a comprehensive range of precision measuring instruments. Our moment weight scales are just one example of our many superior products. For more information about our blade balancing products and services, visit our Blade Balancing  page.

To obtain a price for any of our products, please fill out an online quote request form.

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This paper outlines a method of determining product of inertia by making a series of moment of inertia measurements with the object-oriented in 6 different positions. Product of inertia can then be calculated using formulas which involve the rotation angles of the different fixture positions. Moment of inertia is measured by oscillating the object on a torsion pendulum. Since the object moves very slowly during this measurement, there are negligible centrifugal and windage forces exerted on the object. Furthermore, the mass of the entrapped and entrained air can be compensated for by making a second set of measurements in helium, and extrapolating the data to predict the mass properties in a vacuum.

Click here to download: Using the Moment of Inertia Method to Determine Product of Inertia

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This page describes a new low-speed vertical axis aerospace balance machine that takes advantage of recent advances in technology. This machine measures moment of inertia (MOI) in addition to product of inertia (POI) and center of gravity (CG) offset. Spin speeds as low as 15 RPM yield useful results. This machine has a number of unique features. The operation is totally automatic; even the mass property calculations converting from spin balance to moment of inertia measurement can be accomplished without the operator touching the machine. Gas bearing technology is used throughout, resulting in unrivaled sensitivity and accuracy.

Balancing machines used to require a tedious procedure to adjust the plane separate controls. No matter how many times you did this procedure, it was so complex that you had to have the instruction manual handy as you worked your way through the steps. With the new Space Electronics POI Series Spin Balance Machine, this process is done automatically by the online digital computer that is supplied with the machine. This same computer also prompts the operator with user-friendly menus, so that it is rarely necessary to refer to the instruction manual while balancing the payload.

This machine first rotates the payload at a slow speed and predicts the unbalance forces at the desired speed. If these forces are in excess of the ratings of the machine or the payload, then the operator is warned. He can then balance at a slower speed or stop the machine and determine why the unbalance is so great. Digital filtering techniques are used to reject any forces that do not vary sinusoidally with the rotation speed. If the force transducer outputs include random variations due to air turbulence on the surface of the payload, then the filtering is automatically increased to smooth out this variety.

This paper includes a mathematical mass properties analysis of the errors of measurement as a function of the relative magnitudes of POI and CG unbalance, the moment equations that relate the transducer forces to payload POI and CG offset, and a practical discussion of fixturing and accessory equipment needed to properly balance an aerospace payload.

This instrument is the most accurate slow speed mass properties machine in the world. Unbalance reduction ratios better than 99% are generally achievable, so that the object under test can usually be balanced in a single run. Center of gravity measurement is typically better than 0.001 inch.

The content of this page is derived from a paper presented to the Society of Allied Weight Engineers [SAWE] at Biloxi, Mississippi, by Richard Boynton and Robert Bell of Space Electronics.  Click here to download the complete paper: A New Spin Balance Machine .

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How to measure product of inertia?

Product of inertia can be measured directly by using a vertical two plane spin balance machine (POI series). Spin balance machines spin the payload about a desired axis and detect any unbalances generated. This is by far the most accurate method of determining product of inertia.

But some payloads cannot be spun (some satellites for example), or cannot be spun about all axes (missiles for example). For these payloads, product of inertia can be calculated from moment of inertia measurements (which do not require spinning). Three moment of inertia measurements per plane are required, at widely spaced angles (ideally 0, 45, and 90 degrees). Some moment of inertia measurements can be used in several orthogonal planes, so a total of only 6 measurements is required to get product of inertia in 3 planes. Product of inertia is directly related to these moment of inertia measurements.

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How accurate is this method?

Calculating product of inertia based on moment of inertia measurements is not as accurate a method as directly measuring product of inertia by spinning the payload.

Accuracy of this method depends on several factors, including:

• Accuracy of the moment of inertia measurement instrument used
• Angles used between measurements
• Payload characteristics (payloads with large products of inertia will be easier to measure)

How can Raptor Scientific help me?

Raptor Scientific developed a piece of software that is available with its combined mass properties instruments (KSR Series, MP Series, and POI series). It allows selecting several moment of inertia measurements made in the same plane and angles at which these measurements were made. The control system outputs product of inertia in that plane.

Where do I find more technical information?

For more technical information on this method, we recommend our technical paper Using the “Moment of Inertia Method” to Determine Product of Inertia, which was presented at the 51st annual conference of the Society of Allied Weight Engineers (SAWE).

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When an object is free to rotate, it will rotate around an axis passing through its center of gravity. Therefore it is essential to know moment of inertia through center of gravity to assess the flight characteristics of a payload.

The MOI about an axis A passing through CG is the smallest MOI around any axis parallel to A. Once you know MOI through CG you can extrapolate to MOI through any axis parallel to it providing you know the distance between CG and your axis. The formula is:

MOI through axis A = MOI through CG + Md²

Where:
M is the mass of the object
d is the distance between the CG and the axis A

How to measure MOI through CG without knowing the CG location?

Of course the easiest way to measure moment of inertia through center of gravity is to use an instrument that measures both CG and MOI. Our KSR series are high accuracy instruments that measure CG and MOI with 0.1% accuracy. On this instrument, one payload setup allows to measure two coordinates of center of gravity location and one moment of inertia. The instrument gives moment of inertia results directly through the center of gravity.

Our spin balance machines (POI series) also measure CG and MOI, and therefore give moment of inertia through center of gravity results. Our MP series are instruments that also measure both CG and MOI, although with less accuracy.

But it is not necessary to know the center of gravity position in order to measure moment of inertia through center of gravity. By measuring MOI about multiple parallel axes, one can calculate MOI through CG.

The optimum number of moment of inertia measurements (best compromise between accuracy and time) is 6. More measurements will not provide much more accuracy. Fewer measurements will reduce accuracy significantly.

Moment of inertia accuracy depends on several factors, including:

• The positions used for each measurements
• The accuracy of the instrument
• The accuracy of the fixture

Depending on your payload, you will get an accuracy in the order of 1.5 to 3 times worse than your instrument accuracy. In other words, if your moment of inertia measurement instrument has 0.1% accuracy, you will obtain MOI through CG with 0.15% to 0.3% accuracy.

From one measurement to the next, the payload must be translated in a horizontal plane, without changing its orientation. Moment of inertia measurements give best results when the center of gravity of the payload is located close to the machine centerline. Therefore, measurement positions must be chosen carefully to stay within the tolerance of the instrument for overturning moment, but to provide as great a variation as possible from one position to the next.

Fixturing is crucial for these measurements. The use of a two axis translation table allows measurements without refixturing the payload.

Can I find center of gravity location with this method? What accuracy can I expect?

This method will give you a rough estimate of center of gravity location, but the uncertainty is very large. Do not rely on this method to give you a center of gravity location that you can use in calculation or for balancing purposes.

Accuracy depends on your payload. On a golf club head for example, you can get a center of gravity location to +/- 0.25 inches (+/- 6 mm). On other payloads, the uncertainty can be +/- 1 inch (+/- 2.54 cm) or more.

How does Raptor Scientific implement this method?

Raptor Scientific’s moment of inertia instruments support the multiple MOI measurements method of finding MOI through CG. We design custom software and fixtures that automate the measurement and calculation process.

Our moment of inertia measurement instruments also allow manual center of gravity inputs and translation of MOI results to any axis.

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