Optimum speed rotor

A variable speed helicopter rotor system and method for operating such a system are provided which allow the helicopter rotor to be operated at an optimal angular velocity in revolutions per minute (RPM) minimizing the power required to turn the rotor and thereby resulting in helicopter performance efficiency improvements, reduction in noise, and improvements in rotor, helicopter transmission and engine life. The system and method provide for an increase in helicopter endurance and. The system and method also provide a substantial improvement in helicopter performance during take-off, hover and maneuver.

FIELD OF THE INVENTION 
This invention relates to helicopters and specifically to helicopters 
having variable speed rotors for achieving substantial increases in 
endurance, range, altitude and speed and reductions in noise levels and 
fuel consumption. 
BACKGROUND OF THE INVENTION 
The efficiency of an aircraft, whether fixed wing or rotorcraft, as 
expressed by the fuel consumption required to achieve a specific 
performance as for example, cruise, climb, or maximum speed, is directly 
proportional to the power required to achieve such performance. The power 
required is inversely proportional to the ratio of the aircraft lift to 
the drag (L/D). In order to increase an aircraft efficiency designers 
strive to increase the lift to drag ratio by minimizing the aircraft drag 
at lift levels required to counter the aircraft weight and to allow for 
aircraft maneuvering. 
The lift and drag of an aircraft are determined by the following formulas, 
respectively: 
EQU L=1/2.rho.V.sup.2 SC.sub.L (1) 
EQU D=1/2.rho.V.sup.2 SC.sub.D (2) 
Where .rho. is the air density, V is the air velocity (airspeed), S is the 
reference area of the lifting surface (wing or rotor blade), C.sub.L and 
C.sub.D are non-dimensional lift and drag coefficients. The lift to drag 
ratio L/D is equal to coefficient of lift to the coefficient of drag 
ratio, C.sub.L /C.sub.D. Thus, the ratio of the coefficient of lift to the 
coefficient of drag, C.sub.L /C.sub.D, has a direct effect on performance. 
The C.sub.L /C.sub.D is a function of C.sub.L as can be seen by the 
C.sub.L v. C.sub.L /C.sub.D graph depicted in FIG. 1 for a typical 
airfoil. For best cruise efficiency, the coefficient of lift of the 
lifting airfoil should be maintained at levels of maximum C.sub.L 
/C.sub.D. 
In a helicopter the lift and drag of the rotor blades conform to the same 
lift formula L=1/2.rho.V.sup.2 SC.sub.L where V is the local airspeed on 
the blade which, in a hovering helicopter is a result of the blade angular 
velocity in revolutions per minute (RPM). For convenience, "RPM" as used 
herein refers to rotor angular velocity. Moreover, the term "helicopter" 
as used herein encompasses all types of rotorcraft. 
In a hovering helicopter, the speed of the rotor blade increases radially 
outward. At any given radial distance from the rotor center, the speed of 
the blade is given by the equation: 
##EQU1## 
where, v.sub.t is the rotational speed and r is the radial distance 
measured from the rotor center. 
A helicopter in a substantial forward speed (e.g., 100-200 mph) experiences 
problems of control, vibration and limitations in performance resulting 
from the asymmetry in the speeds of the advancing and retreating blades. 
When traveling in a forward direction 8, the advancing blade 10 has a 
speed equal the rotational speed of the blade plus the forward speed of 
the helicopter, whereas the retreating blade 12 has a speed equal the 
rotational speed of the blade minus the forward speed of the helicopter. 
The speeds along the length of the blades when traveling forward are shown 
in FIG. 2. As a result, the advancing blade has more lift than the 
retreating blade. To avoid helicopter roll over due the airspeed 
asymmetry, the lift on the retreating blade has to be increased while the 
speed on the advancing blade has to be decreased. Because, lift is 
inversely proportional to the velocity (i.e., speed) of the blade squared 
(V.sup.2) a substantial increase in the coefficient of lift (C.sub.L) of 
the retreating blade is required. The available lift coefficient for a 
given blade is limited as shown FIG. 1. Consequently, the asymmetry in 
speeds between the advancing and retreating blades has to be limited 
thereby limiting the forward speed of the helicopter. 
Increasing the RPM of the rotor reduces the relative asymmetry of the 
airspeed distribution, thus reducing the effects of forward speed on roll 
control limits. But such RPM increase is constrained by the maximum 
allowable rotor tip speed. The maximum allowable tip speed is typically 
lower than the speed of sound (i.e., Mach 1) so as to avoid the 
substantial increases in drag, vibration and noise encountered when the 
tip speed approaches Mach 1. 
Current helicopter rotors turn at a constant RPM throughout the flight 
because of the complex and severe rotor dynamics problems. Generally, 
helicopter designers are content if they succeed in the development of a 
single speed rotor, which can go from zero to design RPM when not loaded 
on the ground during start and stop without encountering vibration loads 
which overstress the helicopter and rotor structure. When the blades of a 
conventional rotor are producing lift, a significant change of the rotor 
blade RPM from the design RPM may yield catastrophic results. 
Conventional helicopter rotors are designed to achieve blade flap, lag and 
torsional natural oscillation frequencies, at the operating RPM, which are 
adequately separated from the rotor excitation frequencies occurring at 
the rates of 1 per revolution, 2 per revolution, 3 per revolution and so 
forth. For example, for a rotor operating at 360 RPM, the frequency 
corresponding to the occurrence of a rotor excitation frequency of 1 per 
revolution is 6 Hz (360 RPM is 6 cycles per second), 2 per revolution is 
12 Hz, and so forth. As the rotor RPM is changed so are the excitation 
frequencies. For convenience, the frequencies which give rise to these 
excitation frequencies are referred to herein by the excitation frequency 
occurrence rates. For example a frequency that gives rise to an excitation 
frequency that occurs at a rate of 2 per revolution is referred to herein 
as the "2 per revolution" frequency. For good dynamic behavior, 
considering both blade loads and helicopter vibration, conventional rotors 
with any number of blades are designed to avoid the frequencies of 1 per 
revolution, 2 per revolution, 3 per revolution and so forth. Conventional 
rotor blades are designed to operate at 100% of design RPM with the 
fundamental flap mode at a frequency above the 1 per revolution frequency, 
the fundamental lag mode usually below the 1 per revolution frequency and 
sometimes between the 1 per revolution and the 2 per revolution 
frequencies, and the blade dynamics tuned so that higher flap, lag torsion 
modes avoid the 1, 2, 3, 4, . . . n per revolution frequencies. The 
conventional blade design modes (i.e., modal frequencies) must be kept 
separated from the 1, 2, 3, 4, . . . n per revolution frequencies to avoid 
the generation of vibration loads which may be catastrophic. As a minimum, 
such vibration loads will make the helicopter unacceptable for the pilot 
and passengers and detrimental to the reliability of its mechanisms and 
equipment. To avoid such vibration loads, the rotor angular velocity is 
limited to a narrow range around 100% of design RPM, except for start-up 
and shut-down at low or no rotor load and low wind speed. 
The RPM of helicopter rotors is normally set for a maximum forward speed at 
a maximum weight at a certain critical altitude. The RPM of the rotor is 
such that at maximum forward speed, the tip of the advancing blade is 
traveling at speeds near but below Mach 1, to avoid the substantial 
increases in drag, vibration and noise encountered at speeds approaching 
Mach 1. At any other flight conditions, the rotor RPM and thus, the power 
required to turn the rotor are substantially higher than that required for 
efficient operation. 
Some research helicopters such as the Lockheed XH-51A compound helicopter 
have experimented with rotor RPM reduction at certain flight conditions by 
incorporating a wing for producing most of the required lift and a jet or 
a propeller driving engine for producing the required forward thrust. The 
use of the wings and engine relieve the rotor of its duty to produce lift 
and thrust, thus allowing the unloaded rotor to operate at reduced RPM. In 
this regard, a helicopter can fly at higher speeds before the tip of the 
advancing blade approaches the speed of sound and encounters the increased 
levels of vibration and noise as well as drag. 
Another aircraft, the V-22 Osprey incorporates 2-speed tilt rotors. The 
V-22 Osprey aircraft has wings for generating lift. The rotors are 
typically "tilted" from a first position where their axis of rotation is 
vertical and where the rotor acts as a regular helicopter rotor to a 
second position where their axis of rotation is relatively horizontal and 
the rotor acts as a propeller producing forward thrust. When at their 
first position these rotors operate only to create lift for vertical 
take-off and landing and for hovering. When the needing to go forward, the 
rotors are "tilted" to provide the forward thrust. When in a "tilted" 
position providing forward thrust, the RPM of the rotor can be varied much 
like a variable speed propeller. 
Other attempts have been made in improving helicopter maximum forward 
speeds and/or reducing noise at maximum speed by using 2-speed gearboxes. 
These gearboxes allow the rotor to rotate at two RPM values while 
maintaining a constant engine RPM. The rotor is set to rotate at a lower 
RPM when at high forward speed so as to reduce the rotor tip speed. In all 
other conditions, the rotor is set to rotate at the higher RPM. However, 
these attempts do not substantially improve the efficiency of the 
helicopter by reducing fuel consumption. 
Another helicopter uses 10% reduction in rotor RPM during takeoff and 
landing in order to conform to very strict noise limitations. Because of 
this reduction in rotor RPM, the helicopter performance is compromised 
during take-off and landing. 
While these aforementioned endeavors attempted to increase maximum speed 
and reduce noise during take-off and landing, neither attempted to improve 
the efficiency of the helicopter. Neither attempted to reduce the fuel 
consumed and power required for a given performance or attempted to 
increase a helicopter performance without increasing the fuel consumed and 
the power required. As such there is a need for a helicopter rotor system 
which will improve helicopter range, altitude and speed performance while 
reducing fuel consumption and noise levels. 
SUMMARY OF THE INVENTION 
The present invention provides a variable speed rotor and a method for 
using the same for improving helicopter performance and efficiency while 
reducing fuel consumption. The RPM of the rotor system of the present 
invention can be varied to multiple and even infinite settings depending 
on the helicopter flight conditions to maintain a blade loading for 
optimum performance and fuel efficiency. The present invention allows for 
reduced rotor RPM at reduced forward speeds achieving an increase in rotor 
blade lift coefficient at the lower forward speeds and higher blade lift 
to drag ratio and thus, higher aerodynamic efficiency, lower required 
power, fuel consumption and noise level. By decreasing the RPM of the 
rotor, the power required to drive the rotor at the decreased RPM is also 
decreased. The adjustment to rotor RPM and power can be accomplished 
manually or automatically as for example by computer. 
In order to be able to operate over a wide RPM range, the rotor system of 
the present invention is designed specifically to be able to operate close 
to or on rotor excitation frequencies. To achieve such unique capability, 
the rotor blades are designed to be very stiff and lightweight. The blades 
should be substantially stiffer and lighter than conventional rotor 
blades. 
In an exemplary embodiment, the rotor blades flap, lag and torsion 
stiffness as well as the blade weight per unit length are continuously 
decreasing from the blade root to the blade tip. As a general rule, 
applicant discovered that to achieve operation at a wide range of angular 
velocities, the blades of the present invention require a flap stiffness 
and a blade weight as follows: 
______________________________________ 
Flap Stiffness: 
EI.sub.flap .gtoreq. 25 D.sup.4 
at 10% of rotor radius measured 
from the center of rotor rotation 
EI.sub.flap .gtoreq. 10 D.sup.4 
at 30% of rotor radius measured 
from the center of rotor rotation 
Total Blade Weight: 
W .ltoreq. 0.0015D.sup.3 
______________________________________ 
where D is the rotor diameter and is measured in feet, W is pounds, and EI 
is in lbs-in.sup.2.

DETAILED DESCRIPTION OF THE INVENTION 
The present invention provides for an optimum speed rotor whose RPM can be 
varied to multiple and even infinite settings depending on the helicopter 
flight conditions for optimum flight performance. The optimum speed rotor 
system of the present invention when incorporated on a helicopter allows 
for a substantial improvement in range, altitude and airspeed with less 
fuel consumption and noise levels. For descriptive purposes the optimum 
speed rotor system of the present invention is referred to herein as 
Optimum Speed Rotor or OSR. The OSR can be driven by any powerplant such 
as a reciprocating engine or a turbine engine. 
The present invention allows for reduced rotor RPM at reduced forward 
speeds and/or at reduced rotor lift achieving an increase in rotor blade 
lift coefficient and higher blade lift to drag ratio and thus, higher 
aerodynamic efficiency, lower required power, fuel consumption and noise 
level. The present invention OSR is able to accomplish this while being 
fully loaded, i.e., while producing lift without the aid of a fixed wing. 
Because the lift coefficient of a rotor blade varies along the blade length 
as well as with the blade angular position, it is common to evaluate the 
lift characteristics of a rotor blade by ascertaining its loading. Blade 
loading (C.sub.T /.sigma.) is a parameter which is a function of the rotor 
blade average lift coefficient (C.sub.L) and is defined by the equation: 
##EQU2## 
where T=rotor thrust, S=rotor disc area, V.sub.T rotor tip speed 
T is approximated at T=nW where n is the vertical maneuver factor and W is 
the helicopter weight. The solidity factor, .sigma., is the ratio of 
weighted total blade area to the rotor disc area. 
##EQU3## 
The term "blade loading" or "rotor blade loading" as used herein refers to 
C.sub.T /.sigma.. The useful limit of blade loading for any helicopter 
rotor system can be derived experimentally, i.e., through flight testing. 
The usefull limit blade loading for a typical rotor system is given in 
FIG. 3 by curve 14 as function of the helicopter advance ratio m, i.e., 
the ratio of helicopter forward speed to the rotor tip rotational tip 
speed V.sub.T. As can be seen from FIG. 3 at advance ratios greater than 
0.4, there is a sharp decline of blade loading limit. Thus, to avoid the 
sharp reduction of rotor lift limit, at a maximum forward speed a certain 
minimum rotor RPM has to be maintained to avoid increasing the advance 
ratio beyond 0.4-0.5. An optimum range 16 of blade loading can also be 
derived through flight test for a specific helicopter rotor system as a 
function of advance ratio as shown in FIG. 3. For a given advance ratio, 
the optimum blade loading range is defined by the blade loadings required 
to optimize the various flight performance parameters such as endurance, 
range, and climb rate. 
The OSR of the present invention allows for the adjustment to the rotor RPM 
to maintain a blade loading within the optimum range. By operating below 
100% of RPM, the power required to drive the rotor at the decreased RPM is 
also decreased. The adjustment to rotor RPM and power can be accomplished 
manually or automatically as for example by computer. In a manual OSR 
system, for best endurance, the pilot will manually adjust the rotor RPM 
and engine power to minimize fuel consumption (either directly measured or 
by observing an indication of engine power). For best cruise range, the 
pilot will adjust RPM and airspeed for maximizing the miles traveled per 
unit of fuel. In climb at a given power setting, the pilot will adjust the 
rotor RPM and airspeed for maximizing the climb rate. An automated OSR 
will operate the same way. Information such as fuel consumptions and miles 
traveled per unit of fuel consumed will be monitored by the computer. The 
pilot will select the flight performance parameter that needs to be 
optimized, e.g., range, endurance, rate of climb, etc. and the computer 
will adjust the rotor RPM, power and airspeed settings accordingly for 
maximizing the selected performance. Alternatively, the optimum blade 
loading range as a function of advance ratio is predetermined from flight 
testing and stored on the computer which in turn will adjust the rotor RPM 
and power settings so as to maintain the blade loading within the 
predetermined range for any pilot controlled airspeed and rate of climb. 
Applicant discovered that he can overcome the structural dynamics problems 
associated with significant changes of rotor RPM by building a rotor 
system consisting of blades 18 having reduced mass and increased stiffness 
(FIGS. 4A, 4B, 4C and 7B). The applicant was able to design a blade having 
a continuously decreasing flap, lag and torsion stiffness from the root 20 
to the tip 22 of the blade and having continuously decreasing mass from 
the root to the tip of the blade. The flap 24, lag 26 and torsional 28 
directions are depicted in FIGS. 4B, 4C and 4D, respectively. These blades 
when mounted on a rotor hub will allow for significant changes in rotor 
RPM without being subject to the structural dynamics problems of 
conventional blades. An exemplary embodiment of such a blade is shown in 
FIGS. 4A, 4B and 4C which is made of a carbon-epoxy advanced composite 
material. 
In order to be able to operate over a wide RPM range, the OSR is designed 
specifically to be able to operate close to or on rotor excitation 
frequencies. The OSR is capable of operating a long time under fill rotor 
lift load at or near such frequencies. To achieve such unique capability, 
the OSR rotor blades are designed to be very stiff and lightweight. By 
increasing the stiffness of the blades in flap in relation to the 
feathering axis 30 (FIG. 4A), the blade is better able to operate at or 
near the rotor excitation frequencies. Lag stiffness tends to be less 
sensitive to the excitation frequencies but if kept at a ratio to flap 
stiffness of on average greater than 2 it helps reduce oscillatory lag 
loads and helicopter vibration levels. 
The OSR rotor blades should be substantially stiffer and lighter than 
conventional rotor blades. As a general rule, applicant discovered that to 
achieve operation at a wide range of angular velocities, the OSR blades 
require a flap stiffness and a blade weight as follows: 
______________________________________ 
Flap Stiffness: 
EI.sub.flap .gtoreq. 25 D.sup.4 
at 10% of rotor radius measured 
from the center of rotor rotation 
EI.sub.flap .gtoreq. 10 D.sup.4 
at 30% of rotor radius measured 
from the center of rotor rotation 
Total Blade Weight: 
W .ltoreq. 0.0015D.sup.3 
______________________________________ 
where D is the rotor diameter and is measured in feet, W is pounds, and EI 
is in lbs-in.sup.2. 
The exemplary OSR blade of the present invention shown in FIGS. 4A, 4B and 
4C has a length 32 including the shank 33 of about 17.84 feet, a maximum 
width 34 of about 18 inches and a minimum width 36 at its tip of about 9 
inches (FIG. 4A). The blade has a shank length 40 of about 14 inches and a 
shank diameter 42 of about 3.75 inches. The exemplary blade has the 
dimensions (in.), stiffness (lbs-in.sup.2) and weights per unit length 
(lbs/in.) depicted in table of FIG. 5A. As can be seen from FIG. 5A, the 
exemplary blade has a continuously reducing flap and lag stiffness from 
the hub center to the blade tip. The blade cross-sections at the blade 20% 
station 5C-5C, and the 70% station 5D-5D, are depicted in FIGS. 5C and 5D, 
respectively. The 20% and 70% stations are at 20% and 70% of the rotor 
radius, respectively, as measured from the center of rotor rotation. The 
cross-section of the blade shank is depicted in FIG. 5B. The blade is 
constructed of a carbonepoxy spar/shank and a carbon epoxy leading edge. 
The trailing edge is a lightweight section made of thin carbon-epoxy top 
and bottom skins and a fill-depth honeycomb core. 
In the exemplary OSR blades, adequate torsional stiffness was easily 
achieved. In stiffer OSR blades the use of a hub flexbeams provide for a 
flap and lag effective spring inboard of the feathering axis 30 to reduce 
the load and vibration levels typical of rigid rotor blades. But, the 
spring rate of such flexbeams is not "tuned" to avoid natural 
frequency/RPM "crossings" i.e., the rotor excitation frequencies. 
The blades can be mounted in any type or rotor hub such as hingeless, 
teetering or articulated, to form the rotor system. However, in a 
preferred embodiment, the blades are mounted in a hingeless rotor system. 
A hingeless rotor is well known in the art. It consists of sleeves 59 for 
mounting the blades 18 The sleeves are fixed relative to the hub mast 61 
When mounted on a hingeless rotor hub 60, the blades can not pivot in the 
flap and lag directions relative to the hub (FIG. 6). The preferred 
embodiment hingeless rotor is made of steel. The rotor hub structure is 
chosen to have a hub stiffness in flap and lag matched to the blade 
corresponding stiffness at the blade root. The bearing system 62 
incorporated for blade pitch changes about the feathering axis is also 
required to resist moments that are substantially greater than those for 
an articulated rotor system. 
FIGS. 7A and 7B present two tables, respectively, comparing dimensions and 
design parameters of the exemplary OSR blade incorporated in a hingeless 
rotor to a conventional blade of an articulated rotor system. FIG. 7C 
depicts a scaled comparison between the OSR blade 18 the conventional 
blade 70 being compared in the tables of FIGS. 7A and 7B. The compared 
conventional blade 70 has a length 72 of about 13.17 feet, a constant 
width 74 of about 6.75 inches and a shank length 76 of about 19 inches. As 
can be seen from FIG. 7A, the OSR blades are 85 fold stiffer at about 10% 
radius than the conventional articulated rotor blades which are hinged at 
the root in the flap direction (up-down) and lag direction (forward-aft in 
the plane of the rotor). The conventional blades must be heavy enough in 
order to achieve adequate centrifugal forces to avoid excessive upward 
bending ("coning angle"). In spite of its 85 fold increase in stiffness 
the OSR blades weight per blade surface area is less than half that of the 
conventional blade. This increase in stiffness and reduction in weight per 
blade surface area is achieved on the OSR blades by 3.5 fold increase in 
maximum blade thickness using tapered planform, large root chord and thick 
root airfoils and use of high stiffness/weight carbon-epoxy materials. The 
stiff light weight OSR blades do not require weights at their tips as do 
conventional blades. 
A rotor system of the present invention can operate from 0 to 100% RPM 
under full lift load without reducing the rotor structural integrity. 
Moreover, the vibration levels produced by the rotor of the present 
invention are within acceptable levels as related to crew fatigue, 
passenger comfort and payload performance. The rotor systems of the 
present invention are able to avoid the structural stability, loads and 
vibration problems associated with the operation of the rotor over a wide 
range of RPM. 
The exemplary embodiment OSR blades mounted on a hingeless rotor forminig 
an exemplary OSR were analyzed, optimized and its performance verified 
using 9 integrated dynamics analysis tools for Computational Fluid 
Dynamics, structures, structural dynamics and control dynamics. The most 
important of these tools is CAMRAD II (originated by Wayne Johnson and 
available Analytical Methods Inc, Redmond, Wash.) which was used 
extensively for evaluating rotor stability, loads, vibrations, performance 
and control, including Higher Harmonic Control. All performance and 
structural dynamic data presented are results of CAMRAD II runs with 
non-uniform inflow. In the extensive CAMRAD II analysis, the exemplary OSR 
exhibited no rotor dynamics instability anywhere in the design RPM range. 
The CAMRAD II analysis revealed that the exemplary OSR can reduce its 
angular velocity to as low as 150 RPM (tip Mach number of 0.25) or at any 
other interim RPM to optimize lift/drag ratio, reduce power and achieve 
longer endurance and range or achieve higher altitude and forward speed 
for the same power level. It is expected that the rotor RPM of an OSR can 
be lowered to as much as 40% of the maximum rotor design RPM while 
providing the required lift for a helicopter at its minimum weight. 
FIGS. 8-10 depict power requirements when operating the OSR of the 
exemplary embodiment consisting of 3 blades and a hingeless hub, using a 
low drag unmanned helicopter fuselage, at various RPM values for improved 
efficiency (curve 50) and when operating the same rotor at a constant 
angular velocity of 380 RPM (curve 52), at a helicopter weight of 1400 
lbs, 2600 lbs, and 4000 lbs, respectively at sea level. FIGS. 8-10 were 
created from data obtained from the CAMRAD II analyses. 
The advantage of OSR is dramatic at the lower speeds and light weight range 
(loiter at the end of fuel and with light payload). The reduction of 
60%-70% in power required at 1400 lbs. at 40-80 knots (FIG. 8) provide an 
equal impact on fuel consumption. The reduction in tip Mach number (about 
40%-50%) of the advancing blade may provide 10-15 dBA reduction rotor 
noise levels. The 15 knot increases in speed at constant power of 270 HP 
and the 50 knot increase at a constant power of 120 HP are dramatic and 
indicative of the level of inefficiency of conventional constant RPM 
rotors especially for a low weight helicopter loitering at low speed. 
FIG. 9 shows the performance gains at an average weight of 2600 lbs. As can 
be seen form FIG. 8, the 45% reduction in required power and fuel 
consumption at a loiter speed of about 60 knots will provide an 82% 
increase in maximum endurance for the same total fuel capacity. Similarly, 
the 38% reduction in power required at 80 knots should provide a 61% 
increase in maximum range. 
FIG. 10 shows that even at an overload weight of 4000 lbs. the reduction in 
power of about 25% required at 65-80 knots and the increase in speed at a 
constant power level are substantial. 
A similar power required analysis conducted for hover Out of Ground Effect 
(OGE) indicated that OSR offers 23% increase in take-off weight with 
constant engine power (may provide double the payload weight in most 
helicopters) and 30% reduction in tip speed (may reduce noise level 8 
dBA). The reduction in power required offers 7,000 feet increase in hover 
OGE ceiling out of ground effect with the same engine. 
In an alternate embodiment, instead of operating at a wide range of RPM, 
the OSR can be made to operated at 2 or more angular velocities. With such 
an OSR, the benefits in efficiency will be substantial but not as great as 
the benefits achieved by using an OSR that operates over a wide range of 
RPM.