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Posts posted by iChris

  1. On 4/30/2021 at 6:54 PM, Disguise Delimit said:

    The problem was that the front of the blade, as it rotated, was going down with the main rotor downwash, and was losing efficiency - the back blade was in clearer air, but the loss from the front blade was substantial.

    That’s correct…

     Most U.S. manufacturers will turn the tail rotor clockwise when viewed from the helicopter's left side, taking advantage of the tip vortices coming off the main rotor. When the tail rotor turns in the same direction as the primary rotor vortices, it reduces the relative airspeed of the tail blades, and the available thrust is limited. When the tail rotor turns against the central rotor vortex, the performance increases because of the square-law connection between thrust and increased relative airspeed. 

     Two notable helicopters turn their tail rotor in the so-called wrong direction. They are the MD500 and Robinson R22. However, they both share another less conventional concept, for their time, NACA 63-415 asymmetrical tail rotor blades. More common were symmetrical tail rotor blades like those on the Bell UH-1, AH-1, 204/205/206, 212, 412, and Hughes 269/TH-55.

     Frank Robinson left Bell Helicopter in 1959 and joined the Aircraft Division for Hughes Tool Company, assigned to the U.S. Army's OH-6 Light Observation Helicopter and other Hughes 500 projects. Frank had already established himself as an authority on tail rotor design. Frank found the NACA 63-415 asymmetrical tail rotor blades exhibited noteworthy improvement in performance over the symmetrical blades. 

     Frank brought some of those design characteristics and gave birth to Robinson Helicopters (1973) and the R22. Frank reduced gearing in the tail rotor gearbox to save weight at the aft end of the tail boom. Consequently, due to the engine's reversed position (Front end facing aft), the driveshaft from the engine to tail rotor ended up turning the tail rotor in the wrong direction. Furthermore, asymmetrical tail rotor blades came with an inherent consequence, an undesirable twist or pitching moment. Frank countered the effect with a built-in coning angle designed into the tail rotor.

    The overall compromises ended up cutting weight, and light helicopters like the OH-6 and R22 still provide better than adequate tail rotor performance. MD Helicopters seem satisfied, sticking with the wrong-direction design on their current MD500E/F series. Incidentally, the R44 and R66 tail rotors turn the right way.




  2. On 3/28/2021 at 4:49 PM, sbox23 said:

    @iChris will we be able to know the aux pressure if the pump is operating at a rated speed of 3326rpm and produces a flow rate of 1.5gpm?

    Pressure is equal to force over the area in which the force is applied. The calculation that you’re referring to would only give the theoretical capacity or capability of the pump. You’ll need more configuration or design specifications for an exact pressure value. Begin by taking a look at the pump’s data plate. You need the pump’s horsepower and or torque specification along with what you have.

    Hydraulic Pump Calculations_1

    Hydraulic Pump Calculations_2

    Example taking the pump below:

    Horsepower = ( Q_Flow rate_GPM  x P_Pressure_PSI)  ) / (1714 x Eff )

     hp = (Q x P)/(1714 x Eff)    or   P = (hp x 1714 x Eff) / Q   or 

     P = (torque_inch_pounds  x  rpm) / ( Q x 36.77)  (less 10 -20% for efficiency)

     The Horsepower required to produce 2 GPM @ 1000PSI at 80% efficiency estimates as:

     hp = (2 x 1000) / (1714 x .80)  = 1.46hp


  3. On 4/24/2021 at 5:20 AM, volcanobreath1979 said:

    Can you please amplify? With the main Rotor rotating anti clockwise, the tail rotor should ideally be positioned on the left side of the fuselage to provide anti torque for directional control. On the black hawk, it's on the right (despite main rotors rotating anti clockwise) 

    Numerous factors define the final tail-rotor design, like rotor diameter, tip-speed, blade area, number of blades, blade twist, fin surface area, the direction of rotation, Pusher or Tractor, etc. 

    The least unfavorable compromise is the designer's primary task. Most conventional tail-rotors are Pushers mounted on the left side. Induced velocity below the tail-rotor is higher than above it; therefore, it reduces net thrust if the tail-rotor is blowing toward the fin.

    The exceptions to the Pusher are the Sikorsky UH-60 and Bell, 212, 214, 412. Also, the Bell UH-1, 204, and 205 were initially Pushers, and some were recently converted to Tractors. In 1954 Bell designed the XH-40, the prototype of the UH-1,  AH-1, and subsequent first production Bell 204 and Bell 205 were all Pushers.

     Sikorsky designers chose to tilt the UH-60  tail-rotor shaft so that part of its thrust could help counter CG issues by lifting the helicopter's rear end. The Tractor configuration was chosen to provide clearance with the fin without using a longer driveshaft.

     Bell chose the wrong direction of rotation when in 1954, it designed the XH-40, the prototype of the UH-1 series. Bell's solution 15 years later was to flip the tail-rotor installation from the left side of the fin to the right, using the same hardware. To Bell's good fortune, the tail-rotor blades had no twist, allowing for the change.


    Over the past 45 years, helicopter development teams have discovered that the tail-rotor rotation should have the blade closest to the main rotor going up. A graphic demonstration of wrong and proper rotation was done during the development of the Lockheed Cheyenne attack helicopter. The figure shows the improvement in pedal position obtained when the rotation was reversed. 

    In left sideward flight, the pilot ran out of left pedal between 14 and 18 knots as he tried to fly with the tail-rotor in the vortex-ring state. When the tail rotor's direction was reversed by redesigning the gearbox, the problem disappeared. (The effects of the proximity of the pusher propeller to the main rotor had been discounted by comparing test results with the Pusher on and off.) The change improved even right sideward flight. Just why the tail rotor is sensitive to the direction of rotation in the proximity of the main-rotor wake is not clear.



  4. On 2/13/2021 at 7:48 PM, Agog said:

    I was wondering why the Bell 429 tail rotor is the way it is. I understand the x-shape is to reduce the noise signature, however...

    It is a departure from the typical bell 2-blade tail rotor, so I assume that the 4-blade configuration permits a smaller diameter than an equivalent 2-bladed design.

    My main curiosity is why they chose to go with two 2-bladed teetering tail rotors vs one articulated 4 bladed tail rotor. 


    Bell’s technical description of their 429’s tail rotor follows: “Four blades stacked system, 65” diameter, with low tip speed, scissor arrangement, composite T/R blades with swept blade tips.”

     High blade tip speeds account for significant noise. Noise control can be accomplished by reducing rotor blade tip speed and increasing the number of rotor blades—studies done by members of Airbus, Sikorsky, and US Army referenced below.

    The simplest way to permit flapping is to use a teetering hinge on two-bladed tail rotors. There are simply two teetering rotors spaced a short distance apart on those four-blade tail rotors. Bell stayed with what works. Their time-tested teetering two-blade tail rotor with delta-three.

     Two teetering rotors making up the Hughes/MD 500 four-blade quiet rotor and the AH-64 Apache tail rotor are not at right angles primarily because the scissors configuration simplifies the control linkage arrangement, and there is also some evidence the design is quieter. Note the similarities of three manufactures Hughes/McDonald Douglas,  Bell, and Sikorsky.

     Similarities between the old and the new tail rotor hubs, H269A/AH-64,  Bell206/Bell429, and S58/UH60 below:





    https://apps.dtic.mil/dtic/tr/fulltext/u2/775391.pdf  pg. 115


  5. On 12/12/2020 at 3:07 PM, masum said:

    Looking at performance metrics here, but not clear how much power is used by the tail rotor.   From the power requirements plot therein, look like below 18%

    I realize this may vary by helicopter model.  Anyone heard of a tail rotor consuming > 15% power?  Presumably the higher percentages would be in hover correct?

    You asked, "anyone heard of a tail rotor consuming > 15% power?" absolutely. Flight envelopes account for more than that.

    The numbers are a result of a specific design. There are no metrics other than that from a specific design. Roughly, the tail rotor consumes up to about 10% of the total power for the helicopter. However, allowances of up to 20% may be made for design purposes to ensure sufficient maneuvering and transient capabilities. 

     We can also look at it as a percentage of the total main rotor power. The power required by the tail rotor typically varies between 3% and 5% of the main rotor power in routine flight and up to 20% of the main rotor power at the extremes of the flight envelope. 

     In addition to the yaw function, the flight envelope's extremes require adequate tail power for sideways flight. It is not apparent why anyone would want to fly sideways, but there are plenty of examples. Film cameras are often mounted in the main cabin and only have a clear view of the side. Flying the machine sideways allows the camera to shoot forwards. Flying sideways allows an attack helicopter pilot to dodge fire while keeping his rockets aimed at the target. 

    Pilots regularly fly sideways as a matter of course, because this is what happens when hovering in a side wind. You're flying sideways at the same speed as the wind but in the opposite direction. In the case of a clockwise-from-the-top helicopter, the wind coming from the right side is undesirable as it increases the tail rotor inflow, and so requires more power. The worst-case will then be where the pilot wishes to make a maximum speed yaw-left in a strong wind from the right side. The tail rotor now has to overcome main rotor torque, boom drag due to the side wind. The FAA/military test tail rotor performance under those conditions. The primary criteria to generate enough thrust to balance main-rotor torque in full-power climb with a right-cross-wind with at least a 10% margin left over for directional control.


    Leading University Level Textbook on Helicopter Aerodynamics

     6.9.2 Thrust Requirements

     The primary purpose of the tail rotor is to provide a sideward force on the airframe in a direction and of sufficient magnitude to counter the main rotor torque reaction. The tail rotor also provides the pilot with directional (yaw) control. Roughly, the tail rotor consumes up to about 10% of the total power for the helicopter, although allowances of up to 20% may be made for design purposes to ensure sufficient maneuvering and transient capabilities. 

     This is power that is completely lost, because unless the tail rotor is canted, as on the UH-60 Blackhawk, it provides no useful lifting force. The purpose of the canted tail design is to widen the allowable center of gravity of the helicopter. This design, however. introduces an adverse coupling between yaw and pitch, but this effect can be minimized b) a flight control system. The direction of the anti-torque force depends on the direction of rotation of the main rotor. For a rotor turning in the conventional direction (counterclockwise direction when viewed from above), the tail rotor thrust is to the right (starboard). The magnitude of this thrust, as well as its power consumption, depends on the reaction torque from the main rotor. and the location of the tail rotor from the center of gravity (i.e., the moment arm LMR). In addition, there are inertial effects that the tail rotor must overcome during yawing maneuvers...

     5.4.7 Tail Rotor Power

     The power required by the tail rotor typically varies between 3% and 5% of the main rotor power in normal flight, and up to 20% of the main rotor power at the extremes of the flight envelope. It is calculated in a similar way to the main rotor power, with the thrust required being set equal to the value necessary to balance the main rotor torque reaction on the fuselage. The use of vertical tail surfaces to produce a side force in forward flight can help to reduce the power fraction required for the tail rotor, albeit at the expense of some increase in parasitic and induced drag. If the distance from the main rotor shaft to the tail rotor shaft is XTR, the tail rotor thrust required will be

    Source: Principles of Helicopter Aerodynamics (Cambridge Aerospace Series) 2nd Edition, J. Gordon Leishman (Author) 



  6. On 8/1/2020 at 4:32 PM, cts484 said:

    Because cruise charts are not drag charts, it can be noted the lowest point of a drag chart does not necessarily match the lowest point of the power required curve in a cruise chart.

    1-79. Maximum Endurance airspeed is an airspeed that allows the helicopter to remain flying the most amount of time. It can be found on the power required curve of the cruise chart where power required is at its lowest and not necessarily where total drag is lowest on the drag chart.


    -Why isnt Max Endurance ALWAYS the lowest total drag?  What factors make it not the least amount of drag?

    -What factor(s) make drag charts different from cruise charts/power curves?


    Screen Shot 2020-08-01 at 16.29.22.png

    The text you quoted states that "cruise-charts are not drag-chartsit can be noted the lowest point of a drag chart does not necessarily match the lowest point of the power required curve in a cruise chart."   

     As in Eric Hunt's post above, D = P/V. Were P = rotor power (induced, profile) + the rest of the helicopter (parasitic, tail rotor).

     Eric already answered your question as to why. It's in the math, rearranging the equation D x V = P. It's a helicopter, not just D = P. You  have to account for the V and the other power requirements.

     We're dealing with the total power required supporting more than just the drag of the helicopter. The issue is forward flight (cruising flight) performance. Another power drain is that the turbine engine is more efficient at high power than at low power because of the fuel-flow needed to keep the gas generator spinning, regardless of the power output. Fuel-flow is the center of interest. Remember, fuel-flow is proportional to power; that's why fuel-flow versus airspeed curves mimic the power-required curves. Power is proportional to Fuel-flow.

     To maximize endurance, we want to maximize the amount of time that we can stay in the air. Since the fuel flow is proportional to the power-required, fuel flows lowest when the power-required is a minimum. The speed corresponding to the bottom of the power-required curve is the speed for maximum endurance.

     To maximize the range, we want to get the maximum distance for each pound of fuel burned. Therefore, the maximum range airspeed occurs where a line from the origin is tangent to the power required curve or fuel-flow versus airspeed curve below.


  7. On 7/2/2020 at 1:06 AM, Jones said:

    so  noncommercial helicopter fuel low caution light warning light-on at 70 pounds remaining

    here is my question.  Is it also land soon as possible when 70 pounds remaining warning?

    Refer to the specific series noncommercial/military flight manual. Maybe informational status only indication. Example Army OH-6A manual section below, even though calibration is also 35 pounds:



    • Thanks 1
  8. On 4/24/2020 at 7:10 PM, jake50 said:

    I know my best bet is to ask the FSDO or apply for a legal interpretation, but I'm looking for input on here first. 

    This is mainly referencing 91.213, where it states "small rotorcraft" may operate without an MEL even if an MMEL has be developed for this aircraft.  If you go by small aircraft definition then this would lead you to believe the small rotorcraft is anything under 12,500lbs. 

    I personally wouldn't consider an 12,00 helicopter a small rotorcraft, but I'm no lawyer so I'm looking for input.  When you start reading legal interpretation there are many instances where they somewhat break down rotorcraft into turbine powered or non-turbine powered but never classify them as small or large rotorcraft.  

    In FAA’s eyes a "small rotorcraft:" 

    14 CFR Part 1.1 defines a small aircraft as an aircraft of 12,500 lbs. or less maximum certificated take-off weight.  Therefore, any rotorcraft, could be considered small by the Part 1.1 definition (aircraft) if the rotorcraft/helicopter is less than 12,500 lbs.  

    Part 1.1 Aircraft means a device that is used or intended to be used for flight in the air.
    Part 1.1 Small aircraft means aircraft of 12,500 pounds or less, maximum certificated takeoff weight.

     § 29.811 (f) Each emergency exit, and its means of opening, must be marked on the outside of the rotorcraft. In addition, the following apply: 

    (1) There must be a 2-inch colored band outlining each passenger emergency exit, except small rotorcraft with a maximum weight of 12,500 pounds or less may have a 2-inch colored band outlining each exit release lever or device of passenger emergency exits which are normally used doors.

    • Like 1
  9. On 4/22/2020 at 3:20 AM, Dale said:



    Your speakers are not polarity sensitive so bands 2 & 4 (speaker wires) on the u174 may be reversed wired without a problem. The same for the mic bands 1 & 3 (mic). In most cases the mic is not polarity sensitive. However, with an older mic or special purpose designs, you may have to swap the mic wires around. Once you identify your mic wires, any reverse polarity won’t hurt the mic, it  just won’t work,  just swap it. 

     Your PTT switch above is yellow/black wiring between your radio and the u94. It could be one or two wires. The one wire setup eliminates the extra wire run by using the shield ground at the u94, see photo below. What you have is a momentarily "on" switch that grounds the radio's PTT circuit to key the mic.

    If you’re going to eliminate the switch, you don’t need the yellow wire. If not, the yellow wire should go to one of  the two terminals on the switch and the other switch terminal should have a connection to the black (GND) wire. 

     On one of  the four terminals on the u94 receptacle, you may find a black and white wire soldered to the same terminal. There’re using the black (GND) wire as the speaker return wire

      Isolate-Detect- Correct, the old troubleshooting adage.





  10. From your post, I assume you’re trying to replace the u-94 jack with a u-174 plug or trying to make an adapter cable with a u-174 at each end so as not, destroy the u-94 jack.

    With the documentation at the link below, you should be able the back-track the wiring. Open up the u-94 jack and plug in the u-174 plug. From there, you can back-track the known wiring form the u-174 back to the correct connections on the u-94 side. You can also see how the David Clark H10-76 u-174-plug wiring matches up with the u-94.

    It’s not as hard as it may seem, the system effectively (on the headset end) only uses four (4) wires, two (2) wires to the mic, two (2) wires for the speaker or earpiece. The only reason you have six (6), is they parallel-off two additional wires from the base pair of speaker wires to a second speaker or earpiece. There may be a seventh wire, often used as a shield ground.

    Upstream of the u-94, you normally have six (6) wires. Again, two wires for the speaker/earpieces and two for the mic. The remaining two wires for the Press-To-Talk (PTT). 

    PTT wires are often blue/yellow, mic- red/white, speaker-white/black or white/green. The normally always wires are, red-mic and white or black speaker.     

    Color codes may differ between manufacturers, so don't expect a color-color solution. However, plug and terminal designations are constant.

    See link: Wiring Document U94/U174

    Your u-94 is probably pretty close to one of these  below:

    u-94 color/function
    Red- Microphone High
    White- Microphone Low  
     Speaker High 
    Black- Speaker Low
    Yellow - PTT High  
    Blue- PTT Low 


  11. On 4/10/2020 at 5:46 PM, Spike said:

     if my memory is not failing me, the only time you can switch the grip from flight to idle, in-flight, is during "training" with a "qualified" flight instructor.  Why? First gen B3's had issues with the old style throttle/switch combo which caused a few training accidents while practicing auto's. AIRBUS, then American Eurocopter, was done with it and gave the switch manipulation a limitation.... No more goofin around.....


    On 4/11/2020 at 9:22 AM, Nearly Retired said:

    That's interesting, Spike.  How then does one do a practice autorotation in a B3e?

    It appears your memory hasn't failed you. At least that was the way it was before Airbus.
    Maybe the qualified flight instructor requirement part came in later manuals.




    • Like 1
  12. On 4/8/2020 at 9:39 AM, Discap said:

    I am due to swap out the bleed valve on my Enstrom 480B in the next 25 hours.  In looking at it, it seems to be what I would be termed in the car world as a "waste gate".  I had assumed that it was a control valve but it appears to be simply a preset pressure relief valve.  

    Can anyone share some light on what it does?



    The ability of the compressor to pump air is a function of RPM. At low RPMs, the compressor does not have the same ability to pump air as it does at higher RPMs. To keep the blade angle of attack and air velocity within desired limits and prevent compressor stall, it is necessary to "unload" the compressor in some manner. In other words, the compressor needs to see less restriction to the flow of air through the use of a compressor  bleed air system.




    When the engine is not in operation, the bleed valve is positioned fully open by a spring located inside of the vented piston chamber. The spring along with Pi pressure, directed onto the bleed valve end, are used during engine starting and acceleration to position the bleed valve fully open. 

     During engine operation, Pc pressure is directed through an inlet filter and a restrictor (jet) into the Px chamber. Px air is then directed to Pa through a nozzle (venturi). The rate of air flow from Pc to Px to Pa determines the value of Px pressure for any given N1 RPM. Px pressure is used to provide the closing force on the bleed valve. Px pressure is separated from the Pa pressure and spring chamber by means of the rolling diaphragm. 

     Operation of the valve is a function of preselected ratios of Pi to Pa and Pc to Px to Pa pressures. When Px is less than Pi plus the spring force, the bleed valve is positioned open. When Px is greater than Pi plus the spring force, the bleed valve is positioned closed. The bleed valve is positioned open during engine starting and acceleration until the Px pressure increases sufficiently to overcome the combined value of the spring and Pa pressure. The bleed valve then closes and remains closed at all N1 speeds above the closing RPM.

    Allison Gas Turbine 1981


  13. On 3/18/2020 at 10:39 PM, R22 chopper said:


    I am new here so please forgive me if i get the placing and post wrong. I am very interested in finding out what the legalities would be should one purchase a TIMEX R22 or R44 and fly it under the experimental "banner" or "on condition". My opinion (my opinion only) would be that if one removed the Robinson serial numbers, data plate and anything stating Robinson (for "peace of mind" on Robinsons side in the event of an incident/accident), then would one be able to fly them on condition? I am aware of the Part 91 rule but i am more interested in the Experimental side of it. 

    This would give a second (albeit shorter) life for the TIMEX heli in the case where an owner does not have the cash to rebuild a helicopter, however would like to fly something "safe" (with respect and without stepping on any toes in the current experimental helicopter manufacturer industry).

    You don’t need to rebuild the R22/R44 helicopter or overhaul its engine. However, regardless of the certificate, the aircraft has to be airworthy. It is well-established that an aircraft is deemed 'airworthy' only when it conforms to its type certificate (if and as that certificate has been modified by supplemental type certificates and by Airworthiness Directives), and is in condition for safe operation. Experimental won’t get you pass that. 

    It's a documented practice in line with FAR 43.15c and Appendix D to Part 43. If the aircraft is not used for compensation or hire it could be operated part 91 under the annual inspection only requirements of 91.409a. In that case (with respect to the engine) there would be no required engine overhaul. You could continue on each year as long as the engine passes the annual inspection requirements in Appendix D to Part 43. That’s your on-condition operation.

     Also, as long as the owner complies with chapter 3 page 3.9 or page 3.10 in the R22 maintenance manual, the aircraft and engine can be maintained under FAR 91.409a, 43.15c, and Appendix D to Part 43 in an airworthy condition. To fully understand you may need to read the posts below and the supporting documentation.

      R22 Airworthiness past 2200hrs/12yrs

      R44 12-Year Inspection Required for Part 91?

     Legal Interpretation MacMillan Apr 22, 2011

     FAA Order 8620.2B - Applicability and Enforcement of Manufacturer’s Data 



  14. NTSB Updates on Kobe Bryant Accident

    A ground camera captured an image of the helicopter entering the clouds.

    Radar/ADS-B data indicate the aircraft was climbing southwesterly along a course aligned with Highway 101 just east of the Las Virgenes exit, between Las Virgenes and Lost Hills Road. The helicopter reached an altitude of 2,300 feet msl, approximately 1,500 feet above the highway, but below the surrounding terrain when it began a left turn. Eight seconds later, the aircraft began descending as the left turn continued. The descent rate increased to over 4,000 feet per minute while the ground speed reached 160 knots. The last ADS-B target was received at 1,200 feet msl approximately 400 feet southwest of the accident site.

    A still photo obtained from a security camera located in a road maintenance yard adjacent to Mureau Road and Highway 101 showed the helicopter proceeding westward along the highway and disappearing into the clouds. Mureau runs just to the north of Highway 101. The Board as yet does not know why the pilot entered the clouds. 

    NWS photo looking east from a hill near the crash indicates the tops of the clouds near the site were about 2,400 msl.

    Full text: NTSB Updates on Kobe Bryant Accident By Rob Mark






  15. On 1/31/2020 at 3:36 PM, Jennie said:

     Interesting story- At one time my sister was responsible for deer counts in a city in Iowa.  

    She was asked to fly with heli pilot to do a count.  He told her that the highest cause of death among wildlife biologists was heli crashes.  Whether he was being factual or anecdotal, she politely declined to complete the count.   Thanks for your feedback.  I’m not against this form of transportation. I just think there’s a better solution for this situation.   

    The quote was Aviation accidents....

    Job-related mortality of wildlife workers in the United States, 1937-2000

    “Abstract Wildlife biologists face a variety of job-related hazards that are unique to this profession, most of them involving the remote areas where work is performed and the unusual techniques used to study or manage wildlife. Information on biologists and others killed while conducting wildlife research or management was obtained from state and federal natural resources agencies, solicitations on wildlife-based internet discussion groups, and published obituaries. 

    Ninety-one (91) job-related deaths were documented from 1937 to 2000. Aviation accidents, drowning, car and truck accidents, and murder were the most common causes of death. Thirty-nine (39) aviation accidents accounted for 66% of deaths, with aerodynamic stalls and power-line collisions being the most significant causes of accidents for which information was available. These safety threats should be taken into consideration during the design and planning of future research and management projects.”

     REF: https://www.jstor.org/stable/3784446?seq=1

     Some communities have enacted zoning laws, building codes, fire regulations, etc. that can affect establishment of helicopter landings in residential neighborhoods. 

     They’ve developed codes or ordinances regulating environmental issues such as noise and air pollution. A few localities have enacted specific rules governing the establishment of a heliport. Therefore, contact officials or agencies representing the local zoning board, the fire, police, or sheriff's department, City Council, and the Mayor’s office. 

     Get with your neighbors, kill it at the local level, and the FAA will not approve it in opposition to local laws.  

    Also: http://stophelipad.org/home.shtml


  16. On 1/28/2020 at 1:56 PM, RotorWeed said:

    Anyone have the operating limitations for the S-76b? I think I read once the 76 needs 2 pilots for IFR operations.  

    That S-76B was operated single-pilot VFR, Part 135 charter, limit 9 PAX seats. That wasn’t an IFR operation. That’s also why neither Cockpit Voice Recorder (CVR) or Flight Data Recorder (FDR or Black Box) are required (135.151 or 135.152).


  17. On 1/23/2020 at 12:13 AM, Agog said:


    I'm looking for the single engine fuel burn for a Bell 429.



    22 hours ago, Bonzo828 said:

    It's about 90 gal / hr and can be over 100 if your pulling the guts out....

    The 90gal/hr is a bit high for single engine operation.  However, that's in the range with both engines online, 0.5 - 0.6
     (80 - 97 gal/hr ) in terms of efficiency,  you're looking for < 0.7

    SHP(takeoff) = 1100

    (0.5 * SHP)/6.8  = 80gal/hr

    (0.6 * SHP)/6.8 = 97gal/hr


  18. On 1/23/2020 at 12:13 AM, Agog said:


    I'm looking for the single engine fuel burn for a Bell 429.


    There’s a typical 0.5 - 0.8 Ibs/hp/hr specific fuel consumption (SFC) index for modern turbine engines. Light helicopter turbines, 0.5 is a good average

    The Bell 429’s  one engine inoperative (OEI) 30 minutes hp = 550 SHP. So your OEI would average around:

    40gal/hr. @550 SHP     (jetA est. 6.8lbs/gal)



  19. On 1/2/2020 at 7:49 AM, Maurizio said:

     Governor failure

    Good morning. Can you help me please.
    After the start and heating, ther is a bad situation. the light of Governor do not turn OFF.

    At a certain point the interruptor and light seems working correctly but over the 80% RPM, the Governor do not start.
    It is the control unit?
    Thank You

    Delorean’s post above gives an excellent description of the system. 

    The system is fairly simple. You have a solid-state control unit mounted behind the left seat back. The controller senses engine RPM via tachometer points in the engine right magneto and provides a corrective signal to the governor assembly. 

    The governor assembly gearmotor is attached to the collective stick assembly behind the left seat. When activated by the controller, the gearmotor make the required RPM adjustment by driving a clutch connected to the throttle. 

    The following is a quote from the R22 maintenance manual:

    “The majority of governor problems are caused by the engine's right (helicopter left side) magneto tachometer contact assembly (points) being out of adjustment or faulty.”

    Garbage-in garbage-out, engine right magneto or wiring problems upstream of the controller can result in strange or intermittent issues. The controller may not be the problem.

    Check out the link below pages 8.34A RPM Governor System and 8.34C Governor Troubleshooting


    click photo to enlarge


  20. On 11/16/2019 at 11:10 AM, barrett5991 said:

    Alright everyone,

    Has anyone ever experienced Hydraulic pump cavitation on the R44. I got to thinking about this the other day when I noticed a very faint noise as I tested a hydraulic system. The noise would obviously come on when I turned the pump to the on position, it was a faint whining sound. My only experience with hydraulic pump cavitation has been on large turbine fixed wing aircraft where you would have no chance of hearing any onset failure prior to it happening. The other consideration with this noise would be aeration of some sort. So if you are familiar with this topic, additional knowledge would be appreciated as to get a better understanding of what an impending failure will truly sound like. 

    Cavitation can be recognized by sound. The pump will produce either a whining or a rattling sound. If you hear either, you'll need to determine the source.  These sounds don’t guarantee a hydraulic system problem.

     Flow restrictions, buildup in the strainers, filters, or shutoff valves not fully opening are often causes of cavitation. High oil viscosity, oil that is too viscous and will not flow easily also causes problems. Oil viscosity must be appropriate for the climate and application. Correct any fluid leaks.

     Likely there's not much wrong, the R44 hydraulic system is fairly simple. Maybe just a loose fitting or replacement of the 10-micron filter. You’ll need more evidence of a problem, beyond what you’re hearing.


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  21. On 11/19/2019 at 8:46 AM, Weads said:

    have a simple question on how the inducer bleed port works for the suck/blow hole on the 206. Air escapes the inducer bleed port during startup to allow rapid acceleration and minimize compressor stall of the compressor and it sucks air In to allow for maximum efficiency of the compressor and engine once the engine is at a higher RPM. Although a novice question for most of you guys I can’t figure out how the inducer port is regulated? (I understand it’s a physical hole that’s not actually opened or closed)I know the inducer bleed port is just a slot machined at the top of the shroud housing (compressor lining) but how is air regulated entering and exiting the inducer bleed port. Thanks in advance!

    The 250-C30 compressor bleed air system permits rapid engine response. The system consists of a bleed control valve located on the front face of the scroll and an inducer bleed manifold which encases the slotted compressor shroud housing.

    The inducer bleed system is composed of circumferential slots on the impeller
    inducer shroud, a circumferential collecting plenum, and a bleed port which is located 
    clockwise from the top dead center.

    At low speed, the inducer bleed system bleeds air out, increasing air flow rate at the
    face of the impeller, which reduces inducer angle of attack and decreases the chance of
    inducer stall. This improves part speed stability, especially near 85 % N1, where the
    At high speed, the bleed system sucks in air from outside, which reduces inducer choking. On the standard operating line, bleed direction changes between 95% to 100% N1.

    The compressor receives air at the center of the impeller in an axial direction and accelerates the air outward by centrifugal reaction to its rotational speed. At the lower rotational speed some air bleeds off via the inducer bleed port. As rotational speed increases and the air continues its acceleration, the static pressure decreases according to Bernoulli’s Principle. This continuing reduction is static pressure at the impeller, versus outside ambient pressure, works to eventually choke off the inducer bleed air and allows air flow in through the bleed.  


    During the Allison 250-C30 development program a localized dip in the surge line was encountered. Detailed analysis of the impeller inlet static pressures revealed that the surge line around the 80-85% speed line was influenced by inducer stall. The problem was resolved by adding an inducer bleed and a bleed valve at the compressor discharge that would operate at low speeds. The inducer bleed would bleed air out at low speeds, increasing the airflow into the inducer, reducing blade incidence angles and possibly reducing the boundary layer thickness on the shroud. At high speeds, air flows in through the bleed, reducing inducer choking. The inducer bleed was found to increase the compressor efficiency by 1.5 - 3.2 % between the 85 % and 100% speed lines.

     Ref:  Chapman, Dennis C., Model 250-C30/C28B Development, AGARD-CP-282, 7-9 May 1980, pp. 20-1 to 20-6


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  22. On 7/18/2019 at 4:18 AM, sbox23 said:

    why only the aft vertical shaft and combining transmission auxillary low pressure lights up at 10 PSI whereas the rest of the transmission low pressure lights up at 20psi?


    On 7/23/2019 at 6:15 AM, sbox23 said:

    was really thinking whether both system got to do with the low amount of oil flowing through that why it only require 10 psi.Still hoping that someone has the answer.

    We need to understand the system we’re dealing with. The pump doesn't pump pressure. The pump delivers a rate of flow and that rate of flow meets with and hopefully overcomes resistance in the system. What you’re reading on the gage is not the amount of pressure the pump is putting out. What you’re reading is the amount of resistance being overcome downstream of the gage. Contrary to your quote, 10 PSI represents low resistance and adequate flow rates. The difference between pressure and flow is often misunderstood.

    The CH47 aft transmission lubrication system, see photo below, is a parallel-series system where the pump is servicing multiple branches. The branch could have multiple series loads or additional parallel flow paths. These parallel and series combinations behave differently and have branch pressure that differs from the overall system pressure.

    The transmission pressure is taken downstream of the filter. The filter and transmission resistance to flow causes a pressure drop. The physics of this series branch states the sum of the pressure drop must equal the system pressure. Under normal operations, 6-10 PSI is needed to overcome filter resistance resulting in a 6-10 PSI drop across the filter. The remaining 14-10 PSI is dropped across the transmission.

    Problems with the system similar to filter blockage or blockage downstream of the filter will increase the pressure drop across the filter, and reduce flow into the transmission. The physics are satisfied by an increased pressure drop across the filter and decreased pressure drop across the transmission. The total of all pressure drops in the branch remains equal to system pressure.

    Without the transmission pressure information, we would never have known there was a problem, since the system pressure is 25 PSI (maximum allowed by the relief valve) see photos below.


    click photo to enlarge






    The Difference Between Pressure and Flow

    Pressure Drops in Series Circuits

    Pressure in Parallel Circuits

    Pressure and Flow in a Hydraulic System and Their Basic Relationship



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  23. I have never heard or read an answer to this question that was satisfactory to me. The short answer to this question has been to provide lift to the tail. Two questions should come from this answer. 1 how much of whatever lift is provided is lost in an autorotation? Answer - all of it. 2. Why does this helo need lift in the tail when other single rotor helps don't.


    There’s been much written on the canted tail on the H-60, as evidence in this post; however, you must analyze the tail section dynamics in its entirety (tail rotor, tail pylon, vertical stabilizer, horizontal stabilator, microprocessor controlled stabilator-incidence angle, etc.). Numerous technical papers and article from NASA and others, two listed below.


    “Some basics first: The canted tail rotor is tilted so that some of the rotor thrust is directed upward, which means it contributes to the total lift of the aircraft. The cant angle is 20 degrees, so the tail rotor thrust in the vertical axis is over 30% of its total thrust, while the horizontal axis retains about 94% of the total thrust, a small cost to pay for that lift.


    The two benefits for the H-60 and H-53E are that the lift from the tail rotor help the CG of the aircraft. The 53E third engine was added to the 53D and was placed aft of the transmission, so the tail rotor was used to retain aircraft balance. For the H-60, the aircraft was designed to fit inside a C-130, and so was made low and longer relative to its required payload and volume. The tail rotor helped the designers retain good longitudinal balance.


    For the other aircraft with canted tails, the S-92, the AW-139 and the Bell 525, the lift from the tail rotor helps payload. For the S-92, the canted tail rotor is worth between one and two extra passengers."


    Nick Lappos, Technical Fellow Emeritus


    You can also download Ray W. Prouty's article titled, Evolution of Sikorsky Tails, link below:

    Center of Gravity & Evolution of Sikorsky Tails


    Some try and make sense out of these designs. They often don’t make sense. From a designer’s quote below, it’s just the ‘least worst compromise’


    “Many detailed decisions concerning the rotors and the stabilizing surfaces need to be made before the design is complete. The most sophisticated of the computer preliminary design programs contain logical procedures for making some, but not all of these.


    Many have to be based on considerations that are impossible to computerize and depend on factors ranging from solid scientific fact to controversial aesthetic judgment.


    In almost all cases, there are powerful arguments pulling the designer in opposite directions. Resolving these dilemmas so as to achieve the ‘least worst compromise’ is the designer's primary task.”

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