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I have a question about the tail rotor on a Blackhawk. Why is it tilted the way it is? I have wondered this for a while and I got a look at the Delaware ANG's machine today. I saw it during my lesson today at EVY (Summit, DE). I also thought the tail rotor was on the left side but this one was on the right. I need to go and check photos to see if I am wrong about that one though.

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The tail rotor is on the right. It's canted 20 deg to provide more vertical lift. The design was originally used on the H-53, and Sikorsky carried it over to the H-60.

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I don't remember, I think someone had said due to the excessive aft CG tilting the rotor was a way to help keep the tail up higher. Have you ever seen them do autos? When they flare I have seen some really tail low. Any blackhawk drivers around who can help??????

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I flew the Seahawk for 3 1/2 years at Mayport, FL. The Seahawk and the Blackhawk have the same tailrotor.

The cant may cut down on translating tendency and also limit the effects of various tailrotor effects, like tailrotor vortex ring, etc. I haven't look in the NATOPS (operating manual) in a while, but I think it only talks about the added vertical lift, most likely to correct the CG issue you talked about.

 

The flare in autos is excessive, but I don't think the CG plays into it. It's a 20,000lb helo and there's a lot of momentum. The flare starts at about 200ft and it's not too bad there, maybe 10 deg. At about 100 ft it starts to get big, maybe up to 35 or 40 deg. You can't really see out the front anymore.

 

I've heard that the 53 autos similarly. I've also heard autoing the 53 compared to autoing an apartment building. :lol:

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Tail rotor assembly.

 

(1) Tail rotor is to provide directional and anti-torque control for the aircraft plus a measurable lift of 400 lb.. The assembly incorporates two cross beam tail rotor blades with flexible spars that provide flapping and pitch change. Securing of the tail rotor blades to the tail gear box is accomplished by an inboard and outboard retention plate. Lateral inputs of the tail rotor servo is transmitted to the pitch shaft, pitch beam, pitch links and then to the blades. The 20 degrees cant of the tail rotor assembly provides the following advantages: Low profile, 400 lb.. Lift, greater hover and low speed flight stability, increased rate of climb, allows the nose of the aircraft to be short.

 

 

 

Tail rotor assembly.

 

(1) Tail rotor is to provide directional and anti-torque control for the aircraft plus a measurable lift of 400 lb.. The assembly incorporates two cross beam tail rotor blades with flexible spars that provide flapping and pitch change. Securing of the tail rotor blades to the tail gear box is accomplished by an inboard and outboard retention plate. Lateral inputs of the tail rotor servo is transmitted to the pitch shaft, pitch beam, pitch links and then to the blades. The 20 degrees cant of the tail rotor assembly provides the following advantages: Low profile, 400 lb.. Lift, greater hover and low speed flight stability, increased rate of climb, allows the nose of the aircraft to be short.

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Number one positive effect would be lift..

Canting the tail rotor could "never" impact drifting. The tail rotor counters moment of main rotor. The amount of force needed is given by moment from main rotor and the distance to its axis. This amount of thrust must be delivered to keep the helicopter steady in yaw direction and will also be the amount of thrust creating the drift. To counter drift the main rotor should be tilted.

Tilting the tail rotor upwards will reduce the amount of counteracting thrust but ad lift. Most likely the tail rotor assembly is a bit too heavy and delivers abundant thrust, thus it makes sense to tilt it.

It’s a question of proportion. The main rotor is so big that the tail must be quite long, therefore heavy, to provide for distance between main and tail rotor. It then makes more sense to ad lift aft than to ad weight in front. A redesign, with all costs involved, to make the whole tail lighter, move the COG or even cant the main rotor is most likely a more expensive option than canting the tail rotor..

Edited by ECD

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The tail rotor is Canted 20 degress to provide lift (2.5% of total lift). This in turn improves CG. As far as the Flare at the bottom of an Auto, the Army recommends it at 50 to 75 ft AGL and no more that 25 degrees to prevent stabilator/ground contact.

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The tail rotor is Canted 20 degress to provide lift (2.5% of total lift). This in turn improves CG. As far as the Flare at the bottom of an Auto, the Army recommends it at 50 to 75 ft AGL and no more that 25 degrees to prevent stabilator/ground contact.

 

 

 

 

I couldn't agree more and, a side note the 2.5% lift is at a hover only.

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It's been a a while for me, but if I remember correctly, here's the deal.

Technically, 2.5% lift out of the -10 is only applicable when gross weight is exactly 18,625 lbs. and Nr is exactly 100%.

Multiple replies to the OP are correct in stating that the t/r provides approx. 400-500 lbs of vertical lift. Along with the speed-dependent stabilator, it allows for a a significant enlargement of the safe CG envelope.

The canted t/r increases the complexity of flight control cross coupling considerably. There are dedicated mechanical flight control mixing units for four different aerodynamic coupling effects, which minimize inherent control negative coupling: collective to pitch, collective to yaw, collective to roll, yaw to pitch. The pilot never sees his flight controls move, but inputs to both the m/r and t/r are interdependent and are coordinated mechanically between the trim/SAS motors and the hydraulic servos.

There is an additional electronic coupling that takes place on top of the 4 mentioned above: collective/airspeed to yaw. It helps to compensate for the torque effect caused by changes in collective position. It has the ability to decrease t/r pitch as airspeed increases and the t/r and cambered vertical fin become more efficient. As airspeed decreases the opposite occurs.

Overall, this is a very amazing albeit complex flight control system. And all of this primarily because the t/r is canted 20 degrees....

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Technically, 2.5% lift out of the -10 is only applicable when gross weight is exactly 18,625 lbs. and Nr is exactly 100%.

 

Well the -10 is based on the design weight of 16,825 lbs. The -10 dated 17 April 2006 states on page 72 under paragraph 2.51 "The tail rotor head and blades are installed on the right side of the tail pylon, canted 20° upward. In addition to providing directional control and anti-torque reaction, the tail rotor provides 2.5% of the total lifting force in a hover." After looking through some maintenance manuals and speaking with some MTP's (Maintenance Test Pilots) it provides 400 lbs. of lift and spins at 1,190 RPM's. While 2.5% at design wight is equal to 420.6 lbs. of lift. However, even at heavier gross weights your RPM's are still the same and it will still provide you 2.5% of lift. There may be a thresh hold to this but no definitive answer that I could find. I guess the only way to truly find the thresh hold is to ask Sikorsky.

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It's been a a while for me, but if I remember correctly, here's the deal.

Technically, 2.5% lift out of the -10 is only applicable when gross weight is exactly 18,625 lbs. and Nr is exactly 100%.

Multiple replies to the OP are correct in stating that the t/r provides approx. 400-500 lbs of vertical lift. Along with the speed-dependent stabilator, it allows for a a significant enlargement of the safe CG envelope.

The canted t/r increases the complexity of flight control cross coupling considerably. There are dedicated mechanical flight control mixing units for four different aerodynamic coupling effects, which minimize inherent control negative coupling: collective to pitch, collective to yaw, collective to roll, yaw to pitch. The pilot never sees his flight controls move, but inputs to both the m/r and t/r are interdependent and are coordinated mechanically between the trim/SAS motors and the hydraulic servos.

There is an additional electronic coupling that takes place on top of the 4 mentioned above: collective/airspeed to yaw. It helps to compensate for the torque effect caused by changes in collective position. It has the ability to decrease t/r pitch as airspeed increases and the t/r and cambered vertical fin become more efficient. As airspeed decreases the opposite occurs.

Overall, this is a very amazing albeit complex flight control system. And all of this primarily because the t/r is canted 20 degrees....

 

I thought all of those couplings (collective to pitch, yaw, roll) were for stability in a hover?

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I thought all of those couplings (collective to pitch, yaw, roll) were for stability in a hover?

 

Well what as350b3 stated is correct about the inputs. If you look in TM 11-1520-237-23 (AFCS) on pages 12-5 and 12-6 it will state what he did almost verbatim.

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I thought all of those couplings (collective to pitch, yaw, roll) were for stability in a hover?

 

Yes they primarily provide stability in a hover but it is a mechanical linkage that is still present in flight. In flight, stability is primarily maintained by the pilot (obviously), and the Automatic Flight Control System. As AS350b said it is a complicated flight control system (a proverbial can-of-worms). Even at a hover AFCS is providing assistance to the pilot whether it be heading hold or short term rate dampening (e.g. the effects of wind gusts). To sum it up, stability is a combination of mechanical linkage, pilot input, and electro-mechanical input from the AFCS.

Edited by chamerican

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Is there much difference in basic handling of S-70B, compared to standard S-70 ? I assume the tail folding modifications, and changing tail wheel position, and forward avionics compartments mixed the C.G. a bit, and the main rotor hub is also modified for blades folding. I'm also wondering if the AFCS is also modified because of the time those machines spend in hover...

 

Sorry for those unrelated with topic questions, but I can't find manual for that bird, maybe the NATOPS A1-H60BB-NFM-000 (and whatever JayHawk drivers got) is somewhat restricted, as I was able to find the TM 1-1520-237-10 and TM 1-1520-253-10 with no joy on the old SeaHawk.

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Is there much difference in basic handling of S-70B, compared to standard S-70 ? I assume the tail folding modifications, and changing tail wheel position, and forward avionics compartments mixed the C.G. a bit, and the main rotor hub is also modified for blades folding. I'm also wondering if the AFCS is also modified because of the time those machines spend in hover...

 

Sorry for those unrelated with topic questions, but I can't find manual for that bird, maybe the NATOPS A1-H60BB-NFM-000 (and whatever JayHawk drivers got) is somewhat restricted, as I was able to find the TM 1-1520-237-10 and TM 1-1520-253-10 with no joy on the old SeaHawk.

The mechanical mixing is the same in both models. The AFCS is different. They both have SAS1 and SAS2, although I believe SAS1 may be different between the two. The SH-60B has nothing in the start checklist to ensure SAS1 is off. The UH-60 has Flight Path Stabilization (FPS) while the SH-60B has Autopilot. The latter incorporates an altitude hold (RadAlt and BarAlt), an automatic approach to hover, and a coupled hover. The collective has a trigger to engage and disengage the collective trim on the SH-60B. The UH-60 has a friction lock for the collective.

 

There are alot of other minor differences that make the transition between the two annoying, such as no rotor brake, less sturdy landing gear, no ECS, no built in hydraulic rescue hoist, no torque limits for rotor engagement (due to the tailwheel being further back) on the UH-60, different locations for the parking brake, tailwheel lock, boost pump switches, rotor deice.

 

The UH-60 has a cyclic trigger switch to slew the stabilator up, the SH-60B, with shielded wiring for the stabilator control, does not. The Army also has a bunch of EPs for the UH-60 that are not in the SH-60B NATOPS which include: Lightning Strike, Uncommanded Nose Up/Nose Down, FPS failure, and Pedal Bind.

 

I also noticed this weekend there is no positive open lock on the pilot's door in the UH-60. The SH-60B door must be "unlocked" from the full open position by turning the door latch. The Navy knew to expect gusty winds. I guess the Army figures that you won't leave your hands, feet or anything else near the door opening while sitting in the aircraft. I nearly lost some fingers when a gust of wind pushed the UH-60 door closed with a slam on Saturday.

 

Oh, and who had the bright idea you need a key to unlock and start a military helicopter? The "ignition" switch in the UH-60 is actually like a car, requiring a key. If I look on the overhead console for the igntion switches (#1 and #2) one more time I am going to start banging my head against the glare shield which, as luck would have it, is padded in both models.

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The lift produced by the Blackhawks T/R should be directly proportional to the pedal input, right? Will the nose go up noticably in a right pedal turn?

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The lift produced by the Blackhawks T/R should be directly proportional to the pedal input, right? Will the nose go up noticably in a right pedal turn?

 

 

Having learned a lot since my previous post in this thread...

 

The mechanical mixing unit (MMU) provides yaw to pitch coupling, to compensate for the lift of the tail rotor.

 

So right pedal causes a decreased t/r lift vector, nose pitches up, and the helicopter drifts aft. The m/r tilts forward to compensate.

 

(I cheated, and used one of my books that deals specifically with left pedal inputs, but it's just the opposite for a right pedal turn)

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Oh, and who had the bright idea you need a key to unlock and start a military helicopter? The "ignition" switch in the UH-60 is actually like a car, requiring a key. If I look on the overhead console for the igntion switches (#1 and #2) one more time I am going to start banging my head against the glare shield which, as luck would have it, is padded in both models.

 

The idea for a key in all army aircraft came from an incident in the late 70s, or early 80s where a disgruntled student stole a UH-1 from Fort Rucker an flew it to the White House. At least that's the legend that everybody is told.

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The idea for a key in all army aircraft came from an incident in the late 70s, or early 80s where a disgruntled student stole a UH-1 from Fort Rucker an flew it to the White House. At least that's the legend that everybody is told.

 

A 750 mile trip? And they couldn't stop the guy??

Edited by heli.pilot

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The idea for a key in all army aircraft came from an incident in the late 70s, or early 80s where a disgruntled student stole a UH-1 from Fort Rucker an flew it to the White House. At least that's the legend that everybody is told.

 

it was about 80/81.

it was a digruntled crew cheif that took the bird for a joy ride when the Army in it's infinate wisdom said that crew chief's would no longer have flight status ability. Wally, Gomer or bossman might be able to confirm that.

it happened a couple years before I joined. (84)

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Why was the tail rotor system put on the right hand side of the vertical stabilizer?

 

The commitment to design a tilted tail rotor (2Odeg) demanded that the tail rotor be placed on the right side of the

 

vertical stabilizer or fin. This is called the tractor’ design. and is not the best side for a tail rotor. The induced

 

velocity on the downstream side of the tail rotor is higher than the velocity on the upstream side.

 

For this reason the optimum design is to have the tail rotor sucking air past the fin rather than blowing air at the

 

higher velocity past the fin as in the tractor design. The drag of the fin is higher if the tall rotor is blowing on it

 

which reduces the effective net thrust of the tail rotor. Ray Proutv depicts the two types of tail rotors and the power

 

penalty due to blockage.

 

For example, with the same separation distance between fin and rotor (X/A) of .4 and the same blockage ratio (S/A)

 

of .2, the power ratio penalty is 1.25 for the tractor and 1.1 for the pusher type. This example shows a 15% penalty

 

for the tractor design over the pusher design.

 

A tilted pusher tail rotor could have been designed for the Blackhawk. but the separation of rotor and fm would

 

need to be significant considering tail rotor flapping. The increased separation results in larger components and

 

heavier tail, thus more adverse aft CG effects. There is a penalty with the tractor tail rotor, but the beneficial effects

 

of the tilted rotor resulted in the Blackhawk design.

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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. Answer - this is my educated guess as a seahawk, blackhawk Sikorsky tech rep for over thirty years. The Blackhawk design requirements for payload and transport inside of a C130. The disc loading and weight lifting requirements determine the rotor disc diameter. The required disc diameter determine the rotor blade length. To make the helo fit inside the C130 the rotor center had to be placed forward of the ideal center of gravity to get the long rotor blades in the C130 cargo hold. The lift produced by the tail rotor added much complexity to the mechanical and automatic flight controls. Sad thing was the C130 load requirement was short lived.

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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.”

Edited by iChris

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Everything's a compromise in engineering. If you look at graphs of the sine and cosine you'll notice that as you rotate a bit from zero degrees, you get a lot of sine (vertical thrust) without losing much cosine (horizontal thrust). The vertical is also at a very convenient location, all the way at the tail, which lets you add payload aft of the M/R, in the case of the hawk: fuel, in other designs, passengers or cargo. It costs some weight in terms of structure and complexity, but in order to meet the C130 transport requirement that's what Sikorsky decided they needed to do. They were forced to add the stabilator during flight testing because the M/R couldn't do enough to compensate for the problems induced by the canted T/R.

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