Nomenclature

AEW

airborne early warning

CRP

any contra- or counter-rotating propeller

EPNL

effective perceived noise level

IC

internal-combustion (piston) engine

SPL

sound pressure level

SRP

single rotation propeller

TP

turboprop

rpm

rounds-per-minute

-p

contra-rotating propeller – pusher

-t

contra-rotating propeller – tractor

b

blade span

c

blade chord

${C_P},{C_Q},{C_T}$

power, torque, thrust coefficient

D

propeller diameter

J

advance ratio

M

Mach number

N

number of distinct CRP

${N_b}$

blade count per blade row

P

power

Q

torque

R

propeller radius

T

thrust

V

flight speed

${W_p}$

propeller weight

x

distance between propeller disks

$\eta $

propulsive efficiency

$\sigma $

propeller solidity, ${N_b}c/\pi R$

$\theta $

collective pitch angle

$\Omega $

angular speed

${[.]_{1,2}}$

fore and aft propeller (or blade row)

2.0 Propeller layout

Since at least the 1940s [Reference Gunston2, Reference Bayley-Watson8] there has been a distinction between contra-rotating and coaxial propellers. Blades on one rotating disk are a blade row. A contra-rotating propeller is a unit made of two blade rows arranged to rotate at the same speed in opposite directions. The two blade rows are linked to each other via a common drive shaft. Pitching and feathering is controlled simultaneously, and there is a single propeller governor. Normally, this arrangement is driven by a single engine, such as was the case of the Rolls-Royce Griffon 85 on the Spitfire Mk 22, or by twin engines, via a central gearbox.

Coaxial propellers are separate propeller units mounted on coaxial shafts, each driven by independent engines. This implies that pitching and feathering are independent. In principle, the two angular speeds can be different, one blade row can be stopped, feathered or folded. Experiments on CRP with a blade row locked or windmilling were carried out as early at 1945 [Reference Bartlett9], but there is no evidence that such a contraption was tried in flight.

The term prop-fan was presumably used for the first time in 1971 by Hamilton-Standard to define an unducted turbofan having the benefits of a turboprop/turboshaft engine with one or two rotors [Reference Rosen10], which are either geared or direct-drive. The term open rotor is sometimes used in place of prop-fan. The main difference between a turboprop and a turboshaft is that in the latter case there is only minimal residual thrust from the exhausts; all the propulsive thrust is generated by the propeller blades.

Airplanes with back-to-back propellers which, although seemingly aligned, do not benefit from the close-spacing aerodynamic interference between rotor disks are not CRP. These systems are appropriately called propellers in tandem. A propeller mounted forward of the engine is defined as pusher; a propeller mounted aft is a tractor (or puller). A sketch displaying these configurations is shown in Fig. 1. Over the years, there have been designs with fore- and aft-mounted engines. Aft-mounted pusher propellers were preferable when the wing had large sweep. However, their perceived advantages of un-obstructed air intakes was offset by severe vibrations and the difficulty to divert the exhausts away from the blades. A tractor- versus pusher prop-fan installation analysis (weight, efficiency, operating costs) was published by Pratt & Whitney [Reference Godston and Reynolds11], and it was concluded, belatedly, that “there is no clear winner”.

Figure 1. Selection of CRP configurations.

As an alternative to the CRP, Betz [Reference Betz12] provided a theoretical justification for contra vanes, which had been tested earlier (1933) in the wind tunnel by Lesley [Reference Lesley13]. This was a two-bladed prop with four-bladed contra vanes. The contra vanes can be used to recover at least some of the swirl. It was concluded that efficiency gains could be obtained, depending on the number of vanes, their diameter and their hub diameter. This concept did not win much support. However, CRP operating with the aft blade row locked works as a propeller with contra vanes and has been demonstrated to yield some propulsion benefits [Reference Turnberg and Brown14].

A series of wind tunnel tests were carried out at NACA during the War years; they culminated in a number of reports [Reference Biermann and Hartman15–Reference Gray and Gilman17] exploring aerodynamic effects and comparisons with single-rotation propellers.

2.1 Gearboxes, drive shafts and systems

Gearboxes include reduction gear, contra-rotating gear, and planetary gear for pitch control. Considerable literature exists that shows all embodiments on gearboxes and drive trains, one example of which is shown in Fig. 2, for a Curtiss-Electric propeller: there are at least 16 different gears and 4 shafts with electric motor drive for pitch control. The blade shanks are at the top of the graph. A Curtiss patent [Reference Brady and Chillson4] of that time showed unusual complexity, with hundreds of parts.

Figure 2. Curtiss Electric dual-rotation propeller pitch control mechanism and relative description, Ref. [5].

The propeller governor (and the propeller electronic control in modern propellers) is a speed control system that in its simplest embodiment consists of an rpm-sensing device connected to a pitch-control mechanism by means of hydraulic systems. Mechanical flyweights and springs have been used in the past. Electrical and magnetic control systems are more common today. There is a variety of governors, which are beyond the scope of this contribution.

An important innovation is attributed to Hamilton Standard with their Superhydromatic propellers. They were made of hollow steel, hence lighter, with rapid pitch-change mechanisms (up to 35 degrees/s), deemed essential for high-power manoeuvres, reversing and stopping on the ground [Reference Fey18]. The blades rotated on their axes by individual hydraulic vane motors housed within the open bore of each blade root. The hydraulic fuel supply was independent from the engine. Each blade row had its own pump and oil supply, which was believed to be appropriate, but failed spectacularly in at least one case, the Hughes XF-11.

Hydraulic pitch control was available from de Havilland and Rotol, and it essentially consisted in using engine oil under high pressure to drive a central rack through the propeller axis.

Due to the structural loads at the shaft, their mechanical complexity involving hydraulic and electrical actuators, they are subject to stress and fatigue; for a long time they have been a weak link in the entire propulsion system. Gearbox failure contributed to the cancellation of some programmes (i.e. the Douglas A2D). A parametric analysis of weights, complexity and efficiency is provided in Ref. [Reference Strack, Knip, Weisbrich, Godston and Bradley19].

In some cases, long drive shafts are inevitable, i.e. CRP mounted on swept wings (Tupolev Tu-95, Tu-114), or aft-mounted pusher CRP (Northrop XB-35, Douglas XB-42). In other cases, the design choice was less fortunate i.e. the French Arsenal VB-10 (1944) had two coupled engines [Reference Gaillard20, Reference Green and Swanborough21]: One was placed at the nose and another one behind the pilot. Each engine drove a blade row that made up a contra-rotating propeller that could be driven independently. The front engine drove the aft propeller; the rear engine drove the fore propeller. It used a Latécoère 299A as a test bed, but there is no evidence that it flew. Similar engine architecture was on the Soviet Bohlkhovitinov Spartak and the Japanese Kawasaki Ki.64 (one engine forward and one aft of the cockpit). Long drive shafts were on the Koolhoven F.K.55, Convair XFY-1 Pogo, Lockheed XFV-1 Salmon [Reference Leonard22], and Tupolev Tu-91 (all engines behind the cockpit).

Due to their complexity, modern propellers require their own certifications; the software control units must comply with the most stringent regulations. Achieving a new type certificate is a major undertaking that is often under estimated. An advisory note from the FAA [23] describes an acceptable method that may be used to demonstrate compliance to the propeller type certification requirements of Title 14, Code of Federal Regulations, Part 35. However, these regulations to do not apply to CRP, as stated in Section 2.1.6 of said document, where it is reported that “The complexity of a contra-rotating propeller has not been considered; therefore, additional requirements may need to be established.” Terms of Reference published by the EASA [24], and agreed by the industry, highlight the difficulties of these new propulsors, and advise that they should be certified under engine regulations, with additional provisions to account for the characteristics of the open rotor. These provisions mention at least uncontained rotor failure and bird ingestion. No CRP is currently certified. Any remaining Russian CRP may follow domestic military certification.

2.3 Propeller parameters

It is important to clarify how the blades are counted, to avoid double counting. Each blade row has its own blade count, say ${N_1}$ and ${N_2}$ . Therefore, the total number of blades is ${N_1} \times {N_2}$ . The confusion stems from the fact that early propulsion systems had the same number of blades, the same diameter and rotor solidity; however this is no longer the case. Thus, we define a diameter ratio ${D_1}/{D_2}$ , a power split ${P_1}/{P_2}$ , a torque split ${Q_1}/{Q_2}$ , an angular speed ratio ${\Omega _1}/{\Omega _2}$ , etc. Normalised axial separation $x/D$ ranges from 0.1 to 0.2 for most CRP, except prop-fans, whose axial separations can be as large as 0.30, with hub-to-tip diameter ratios of 0.35–0.40.

The rotor solidity, $\sigma = {N_b}c/\pi R$ , is calculated for each blade row, although there can be cases where the total number of blades is used if the blade rows are geometrically similar. A parameter that is sometimes used is the activity factor, which is a measure of the geometric width of a blade, and hence the load. The activity factor is calculated from the radial integration of the normalised chord ${(c/D)^3}$ via a multiplicating factor leading to values $\sim\!{10^2}$ ; by contrast, we have $\sigma \lt $ 0.5. It is not possible to evaluate the activity factor without detailed blade data. However, since it is proportional to the blade solidity, the latter parameter is used. Modern prop-fans have complex geometries with variable chord ( $c/D \sim 0.15$ ) and short spans ( $b/D \sim $ 0.4–0.5).

It is common to have differential pitch settings of fore and aft blade rows. These settings have to be finely tuned in order to avoid vibrations, reduce noise and maximise propulsive efficiency. One example cited is that of the Supermarine SeaFire F.47 Rotol CRP [Reference Bailey-Watson30], where the fore propeller was set at 20 degrees and the aft propeller at 23.5 degrees [Reference Fey31].

Other parameters include the net-thrust/disk area ( $T/A$ ) or net-thrust/blade-area ( $T/\sigma A$ ). Loading is also given as shaft-power/disk-area ( $P/A$ , measured in kW/m $^2$ ). The $P/A$ ratio is a design constraint.

Values of $P/A$ for modern prop-fans reach 400 kW/m $^2$ , but exceed this limit in laboratory test models. A statistical analysis of this parameter is shown in Fig. 3. The data points are only approximate, and they rely on estimated power output at take-off, the mean diameter between the blade rows and equal distribution of power; however, this is not always correct. For counter-rotating propellers, it is assumed that each engine powers one blade row. For contra-rotating propellers where a single engine powers both rows, then the power is split equally. Aside from these caveats, there is no doubt that the blade loading has increased considerably over the years. The FH385/-6 with the Europrop TP400-D6 is a single-rotation propeller [32], but it has been added to the chart to demonstrate that highly loaded advanced propellers and prop-fans can provide as much power as a CRP. A graphic comparison among different propellers is sketched in Fig. 4.

Figure 3. Summary of CRP power loading.

Figure 4. Selection of CRP and SRP, including the Ratier-Figeac FH385/-6 (Airbus A400M), the Hamilton-Sundstrand NP2000 (Lockheed C-130H and P-3, Northrop Grumman E-2) and the Safran Open Rotor.

2.5 Engine layout

CRP are powered by single or twin-engines. In the latter case, we define the engines as ‘coupled’. According to Ref. [Reference Leonard22], coupled engines ‘include any of the ways that two engines or parts of engines are put together to create a more powerful engine to jointly drive one propeller or two coaxial propellers’.

Coupled engines are standard propulsion architecture on twin-engined helicopters. For fixed-wing CRP aircraft, known cases include the following: (1) one engine driving fixed-pitch CRP (these are the oldest examples in aviation history); (2) one engine driving variable-pitch constant-speed CRP with blade pitch linked to each other (i.e. Avro Shackleton); (3) twin engines each driving its own fixed-pitch variable-speed propeller (Macchi-Castoldi MC-72); (4) twin-engines driving coaxial variable-pitch constant-speed propellers. In the latter case, there can be two different mechanical solutions: (a) blade pitch control has only one degree of freedom, with the two propellers slaved to each other (Saunders-Roe SR-45); (b) each blade row has an independent pitch control, leading to two degrees of freedom and better power split (Bristol Type 167 Brabazon). An example of twin-engine configurations is displayed in Fig. 5. In case (a), the propeller rpm is the same and the two blade rows are synchronised, and there is one propeller governor (i.e. Convair XFY-1 and Northrop XP-56). In case (b) both pitch and rpm are independent, and the propellers are not necessarily synchronised (Douglas XB-42 and the Fairey Gannet). V-drive cases are known to exist at least for one airplane, the Bristol Brabazon. The default architecture of a twin-engine helicopter has a central shaft as in Fig. 4, regardless of whether it is a single or coaxial rotor, with sprag clutches ${G_1}$ and ${G_2}$ required to disengage one engine if inoperative. There is at least one example of three coupled engines designed to drive a dual propeller: the Allison XT44, with the possibility of disengaging at least one of them [Reference Leonard22]. Several derivatives of the Allison turboprop T40 were used with CRP: Douglas A2D, Convair XP5Y-1 and R3Y-1, Hiller X-18 and North American XA2J.

Figure 5. Examples of twin-engine CRP systems; “G” denotes a gearbox.

In response to a UK Ministry of Supply requirement for a carrier-based airplane, Fairey Aviation developed a coupled engine in the 1930s, the P. 24, Fig. 6(a). This engine was made of two banks of internal combustion engines (750 kW each) each driving a co-axial propeller. The two propeller shafts made up a coaxial variable-pitch 3 $ \times $ 3 CRP developed by Fairey Aviation. When viewed from the front, the fore propeller rotated clockwise powered by the right crankshaft; the aft propeller rotated anti-clockwise powered by the left crankshaft. This engine never entered productionFootnote 1, although it was tested on the Fairey Battle (registration K9370). Its development is described in Ref. [Reference Hollis Williams42].

Figure 6. Four British engines with CRP combinations: Fairey Aviation coupled piston engine P. 24 driving Fairey CRP [[Reference Hollis Williams42]; de Havilland hydromatic CRP with Armstrong-Siddeley Python engines [43]; Napier Nomad I with Rotol CRP (Royal Aeronautical Society Archive); Napier Coupled Naiad with unknown CRP (IMechE Photo Archive NAP /4/3/9/5).

The Napier Nomad I (1948–1952), displayed in Fig. 6(c), was a sophisticated engine that had a turbine driving the fore propeller and the diesel engine driving the aft propeller. The two blade rows were not linked together. They could not be powered independently, because the piston engine was required to power the aft propeller, and the engine’s exhausts powered the turbine, which was needed to power the fore propeller. A second engine version, the Nomad II (1952–1955), mounted single-rotation propellers only [Reference Gunston2, Reference Sammons and Chatterton44]. Evolution from the former to the latter is not explained in the available papers, i.e. Ref. [Reference Chatterton45]. However, Napier recognised its complexity [Reference Howard46] and offered the Nomad II for the Avro Shackleton upgrade (four-bladed SRP, Rotol or de Havilland). Diesel fuel and wide-cut kerosene could be used.

A related engine that did not reach production was the Napier Coupled Naiad, planned for a Blackburn aircraft. At least one photo with CRP is known (1948), shown in Fig. 6(d). In the discussion of Ref. [Reference Hollis Williams42], “the advantage of that engine was that it was a truly contra rotating engine”. It was lighter, powerful and had the gearbox aligned with the turbines. Nevertheless, it was cancelled, due to engine development problems. The Python engine was coupled to Rotol propellers, not de Havilland as shown in Fig. 6(b).

After the end of the War, Fairey Aviation asked Armstrong-Siddeley to develop an equivalent coupled engine architecture based on the gas turbine. The Armstrong-Siddeley Double Mamba (1949–1979) was a coupled engine first designed for the Fairey Gannet [Reference Hollis Williams42], but then used for the Blackburn B-88 and for the Short S.B.3 Sturgeon. Each Mamba drove its own propeller, was completely independent from the other one, had separate fuel feed lines, lubrication and control systems. This engine is equivalent to the sketch in Fig. 5(b), and is also shown in Fig. 7(a).

Figure 7. The British Double-Mamba turboprop (Flight International Archive), and its American competitor, the Allison T40, on the final assembly bench (Allison Archive).

The gas turbine had a 0.0874:1 gear ratio, an 11-stage axial compressor with 5.75:1 compression ratio, and a 3-stage axial turbine. The gearbox and the drive systems weighed 360 kg. Detailed technical data, including power output charts, are reported in Ref. [37]. The starboard engine drove the fore blade row and the port side engine drove the aft blade row. This architecture allowed the propeller to function both as CRP and SRP, with one blade row feathered and braked [Reference Turnberg and Brown14]. Compressed air was used to start one engine; engine exhausts fired the second engine. This capability was exploited to limit fuel consumption, particularly in ground manoeuvre and long-range reconnaissance. Unaware of this precedent, the author co-investigated [Reference Smith, Filippone and Bojdo47] the benefits of using stopped blade rows to reduce fuel consumption and noise emissions at some conditions. Acoustic benefits were demonstrated for blade counts other than ${N_b} = 3 \times 3$ .

2.7 Propulsive efficiency

Early-days concepts focussed exclusively on increasing aircraft speed [Reference McCoy25]. If only propeller thrust were involved, increasing net thrust could only be delivered at higher collective pitch, which would load the propeller to a point of diminishing propulsive efficiency. Thus, the idea was to couple a contra-rotating co-axial blade row to take advantage of the swirl of the former propeller wake, and share the load between two rotating disks. An efficiency gain of up to 8% is sometimes reported [Reference Hager and Vrabel50]. This gain is often mis-quoted, which is somewhat concerning. The angular momentum of the wake depends on the torque imparted to the rotor and on the blade count. Lightly loaded propellers do not benefit from a swirl recovery system. Only propellers at the top end of their design load generate enough angular momentum that could, in principle, be recovered [Reference Mikkelson, Mitchel and Bober51]. Even then, the contra-vanes pitch must be adjustable.

The technical literature reports graphs [Reference Kuznetsov52, Reference Awker53] of propulsive efficiency of SRP, CRP and turbofan engines versus Mach numbers. Those charts are markedly different among various sources; their origin is unclear.

Normally, the best efficiency is estimated at about 80% ( $ \pm $ 2%) for a single-rotation propeller, and 2–6% above these values for a CRP (84–86%), depending on flight Mach number and propeller advance ratio. A notional propeller efficiency plot in terms of advance ratio $J = V/\Omega R$ is displayed in Fig. 8. A theoretical analysis of propeller efficiency [Reference Kroo54] for various propeller-wing configurations indicates that contra-rotating prop-fans are better than single rotation propellers because of swirl recovery, although the wing can contribute to this recovery. The propeller-wing integration must be designed together, with the vertical location of the propeller being critical in the overall performance. For a modern design, we have numerical tools that allow us to make parametric analysis and multi-point design.

Figure 8. Notional propulsive efficiency of a conventional propeller and a contra-rotating propeller.

Alongside propeller efficiency, we need to consider the aero-engine’s thermal efficiency. When comparing two propulsion systems, all the intermediate steps must be accounted for. We have the thermal efficiency of the gas turbine ${\eta _t}$ (or internal combustion engine), the efficiency of the reduction gearbox and all other gearboxes ${\eta _g}$ and the efficiency of the propeller ${\eta _p}$ . The result is $\eta = {\eta _t}{\eta _g}{\eta _p}$ . A complication arises then the turboprop delivers a non-negligible thrust from the exhaust gases, or when we compare different flight regimes. High propeller efficiency is of little help if the turboprop engine has poor thermal performance, or the gearbox produces excessive power losses, or else there are weight penalties.

Propellers are designed to deliver a torque at prescribed flight speeds and altitudes. If the torque available is considerably larger than the design propeller torque, there is a risk of stalling the propeller blades. The design torque can be increased in a number of ways, i.e. by increasing solidity (number of blades and/or blade chord), angular speeds and diameters. Larger angular speeds and diameters inevitably increase tip Mach numbers and thus limit the flight speed. The increase in propeller solidity has the effect of increasing the profile drag, and thus reducing propulsive efficiency. In this instance, the clever idea seemed to be coupling another blade row on a coaxial shaft, and thence produce a contra-rotating propeller with far larger design torque. This solution also required a minimal increase in mass.

Now consider a notional single-rotation propeller with performance chart as shown in Fig. 9(a). This propeller was calculated with the numerical methods explained in Ref. [Reference Filippone55]. An operation point at ${C_Q} = 0.003$ would provide a ${C_T} = 0.0083$ with a propulsive efficiency $\eta = 0.85$ and a speed of 139 KTAS. Doubling of the available torque to 0.006 would ideally almost double the net thrust. This can be achieved at point B, where the efficiency has now decreased to $\eta = 0.75$ , and the flight speed has decreased to 120 KTAS. The corresponding collective pitch has increased. With a fixed-pitch propeller, there is a sharp decrease in propulsive efficiency as soon as the propeller moves from its design point, Fig. 9(b). For example, at $\theta = 30$ degrees, an increase in power is achieved at a loss of 12–14% efficiency (line from C to D): The propeller spins faster and the advance ratio decreases; fixed-pitch propellers are not very efficient.

Figure 9. Notional single-propeller performance, showing the effects of torque, collective pitch, propulsive efficiency and flight speed on the net thrust. Variable and fixed pitch are considered.

Propulsive performance is sometimes quoted as specific fuel consumption (SFC), which is incorrect: This refers to the engine only for power delivered at the shaft, without gearing effects. Data are supplied from rated power and fuel flow, which are not always reliable, and in any case suffer from installation, ageing and maintenance problems. In any case, we find SFC for the NK-12MA turboprop at 0.219 kg/kW/h, against 0.228 kg/kW/h for the TP400-D6 turboprop. These are among the best in this class of engines.

3.0 Pioneering contributions (before 1940)

Frederick Lanchester filed a Patent in 1907 (UK Patent 9413A) and then proposed a propulsion theory [Reference Lanchester56]. His ideas were subsequently revised and appeared in 1941 in Flight Magazine [Reference Lanchester57]. Said invention required the torque to be balanced and opposite, something we now understand is not required. There is no record of this contraption having been built.

It is unclear when the first CRP flew successfully, although there is some evidence of one-off projects in the 1920s and 1930s, and earlier historical photographs showing concepts that had no chance of being successful, i.e. Ferdinand Feber and Roger Ravaud in France. Documentation from that time is rather limited. For example, a report in Popular Mechanics (November 1931) tells of a successful flight of “propellers revolving in opposite direction”, with the implication that one propeller was spinning at half the speed of the other one. A Dutch design, the Dekker-Fokker C.1, registered as PH-APL, mounted short turbine-like counter-rotating propellers running at different speeds. Only taxi trials are known (circa 1937), but it is unlikely that the airplane was cleared to fly. A description of this contraption is available in a US Patent by Adriaan Dekker [Reference Dekker58].

There are verifiable historical records, such as the 1933 flights of the Italian Macchi-Castoldi MC-72 seaplane for the Schneider Trophy, shown in Fig. 10. The airplane had two-bladed contra-rotating propellers ( ${N_b} = $ 2 $ \times $ 2; 2.5 m diameter). This airplane had two Fiat AS.6 IC engines mounted in tandem [59, Reference Bona60], driving two fixed-pitch co-axial propellers. It achieved a speed of 383 KTAS on 23 October 1934, in spite of the drag caused by two large floats Footnote 2.

Figure 10. Sketch of the Macchi-Castoldi MC-72 world-record sea-plane.

There are prototypes that flew once for a few minutes but then were grounded and the programme cancelled, without a definite assessment of its benefits or drawbacks. One example is the Dutch Koolhoven F.K.55 (1938), with 2 $ \times $ 2 Ratier metal CRP (variable or fixed pitch, depending on archival sources). It only achieved a 2-minute flight and suffered serious stability issues. The pilot’s seat was between the propeller gearbox and the engine. This airplane was displayed at the Paris Salon de l’Aviation in 1938 and destroyed by German bombing in May 1940.

Examples of tandem configuration include the following: under-the-wing propellers of the Fokker F-32 (1929) and the Farman F-222 (1932–1940); over-the-wing propellers of the Dornier Do-26 seaplane (1938–1940) and Latécoère 300 in France. These were coupled propellers with one tractor prop aligned with a corresponding pusher, shown in Fig. 1(d). In the Farman F-222/2 case, the distance between propellers was $ \sim $ 2 diameters, with two separate engines on the same nacelle. There are still wake effects, though not in the measure normally found in a closely spaced CRP. Dornier was responsible for similar tandem configurations, with the massive 12-engined Dornier Do-X (1929–1932), with six propellers in tandem (Fig. 11), and the Dornier J Wal (1925–1950), having one tandem four-bladed propeller system.

Figure 11. Sketch of the Dornier Do-X seaplane (1929) and the Saunders-Roe SR-45 Princess (1952).

According to Dornier himself [Reference Dornier61], for the Do-X the best solution with the knowledge at the time was to couple two engines on a single nacelle, with each engine powering its own propeller. Driving one propeller with two engines would have required a too large diameter. Other options would have required heavy gears or long shafts. The four-bladed props were manufactured together into a single piece, making them fixed-pitch wooden props with presumably poor aerodynamic aerofoil sections. Propeller characteristics are not mentioned in the official reports, i.e. Ref. [62].

The use of back-to-back propellers does offer some propulsion advantage without the complications of closely spaced rotors. The aft propeller operates in the wake of the fore propeller; this wake is an accelerated swirling flow. The aft propeller takes advantage of higher apparent speed $ \sim \!V + \sqrt {{T_1}/2\rho {A_1}} $ and recovers the angular momentum by contra-rotation. One claimed feature of this arrangement was that the failure of one engine would not change the thrust line of the remaining engine.

4.0 A golden era for the contra-rotating props (1940–1950)

A great deal of creativity and risk was taken during the War years. During this time, there was a mix of piston engines and turboprops. Piston engines turned out be obsolete in the rapidly evolving world of aviation, with turboprop engines developing rapidly for high speed flight [Reference Holbrook and Rosen63].

The 1930s were dominated by seaplanes with propellers in tandem. They too became obsolete with the development of ground infrastructure that permitted the operation of land-based airplanes. The War years saw an interest in highly powered airplanes, mostly single-propellers capable of operating from/to aircraft carrier: They needed short take-off and landing performance, which was provided by highly loaded CRP. In this instance, the CRP had the benefit of neutralising the torque generated by a conventional propeller, which could have caused the aircraft to roll on its side during a high-power take-off. Just after World War II, a number of military airplanes with CRP appeared and made use of the nascent gas turbine engines, although the transition was slow. These airplanes include the Blackburn B-54, Westland Wyvern, Douglas A2D, Martin-Baker MB-5, Shorts S-38 and others, some of which are described as follows.

In the United Kingdom, the Blackburn B-54 (Blackburn Aircraft) was a carrier-borne anti-submarine aircraft making use of pre-war piston engines, the Rolls-Royce Griffon 56, shown in Fig. 12. It was originally designed around the Napier Double Naiad turboprop (see Fig. 6(d)). A later variant (registered as WB797) using the Armstrong-Siddeley Double Mamba turboprop engine, had $4 \times 4$ CRP. In this form, the Blackburn Type number was renamed to the Blackburn B-88 [Reference Jackson64]. This programme was cancelled when it became clear that the Fairey Gannet was a superior aircraft. The same Double Mamba engine was installed on the Short S.B.3 with two 3 $ \times $ 3 Rotol CRP. The airplane registered WF632 was demonstrated at the 1950 Farnborough Air Show. The looks of this aircraft were certainly unimpressive. Other versions were powered by the Rolls-Royce Merlin 140 engines. In any case, a few of these aircraft were operational until at least 1954, both as carrier and land based.

Figure 12. The Blackburn B-54, WB781 (1949) with Rolls-Royce Griffon engine. It had cranked wings (anhedral/dihedral) that could not fold. Photograph from the Royal Aeronautical Society Archive.

The Westland Wyvern Footnote 3 W.34 (1946–1958) had $4 \times 4$ CRP powered by an Armstrong-Siddeley Python turboprop engine (version TF.4, later renamed S.4), although the earliest version of the airplane, TF.1 shown in Fig. 13, had piston engines (Rolls-Royce Eagle). Its development was delayed by problems in the propeller mechanical design, a Rotol Type 5. The de Havilland product brochure [43] indicates that they planned to have one of their CRP fitted with the Python engine, shown in Fig. 6(b). Its first flight took place in December 1946 and operation was delayed to the 1950s, with the airplane demonstrated at the Farnborough Air Show in 1953; it saw military service in the 1956 Suez Canal crisis. The Wyvern was the last Westland fixed-wing aircraft before the company moved into rotorcraft. At least 127 Wyvern were produced in all versions. One version that was to be fitted with the Napier Nomad I turbo-compound engine was not completed.

In the same vein, in the US, the Douglas company developed the A2D attack aircraft, of which only flying prototypes and pre-production aircraft were available. The programme was cancelled due to several technical problems, presumably due to the complexity of the propulsion system and the Allison XT40 engine problems. The CRP was a single ${N_b} = 3 \times $ 3 pusher configuration, with fixed-rpm fully feathering blades having sharp-cut tips. The rotor solidity was $ \sim $ 0.193.

Figure 13. The Westland Wyvern TF.1, first prototype, registered TS371, with Rolls-Royce Eagle engine. Photograph in the public domain.

In the 1940s, Northrop developed the XB-35 flying wing and thus faced a complex integration issue between aft-mounted tractor CRP and the rudderless aircraft. Most of the documentation available is concerned with the flight stability of the flying wing, i.e. Sears [Reference Sears65]. The engines of the initial version were Pratt & Whitney R-4360 Wasp air-cooled radial internal combustion engines and powered the propellers through a very long drive shaft running through the wing ( $ \sim $ 7 m), as displayed in Fig. 14. Hamilton-Standard developed the contra-rotating tractor propellers (Hamilton Standard type HSP24F60-344). This aircraft evolved into the YB-49 with the replacement of aft-mounted CRP with Allison J35 turbojets. A design with conversion to the Northrop T37 turboprop was named EB-35, but still retained drive shafts across the wing. Even that programme was cancelled, involving industrial relations and government pressure. The programme matured too late (1946), although was intended for war service. The propeller issues were never resolved and involved failures at all levels, from feathering/pitching to vibrations compounded by the long drive shafts.

Figure 14. Sketch of the Northrop XB-35 flying wing and the turboshaft upgrade proposed for the EB-35.

The Convair XP5Y-1/R3Y-1 and -2 seaplanes were novel designs that demonstrated a relative success. The latter one, named Tradewind, had a prominent nose radome, shown in Fig. 15. The airplane was a cantilever high-wing flying boat with four Aeroproduct CRP, ${N_b} = 3 \times 3$ , powered by Allison T40-A-10 turboprops, with gear-ratio 0.0630–0.0638 [66, 67], inter-propeller separation equal to 0.13 diameters and propeller water clearance $ \sim $ 2.7 m. The programme was finally cancelled in 1958 because of engine problems and the loss of three aircraft. The R3Y-2 managed to perform unique feats such as in-flight refuelling of four aircraft simultaneously. The XP5Y-1 had Hamilton Standard HS2F117B-24 propellers, with 4.60 m diameter [68].

Figure 15. Sketch of the Convair R3Y-2 Tradewind mounting Aeroproducts 4 $ \times $ 4 CRP.

The Fisher XP-75A was a costly and ill-fated project [Reference Norton69]. This airplane went into pre-production (about 14 built or half built). Of interest was the fact that two types of CRP were tested. Initially, a 3.96 m diameter, narrow, A-blade model was used. However, some of these aircraft flew with another propeller model, a ${N_b} = 3 \times 3$ blades, with 3.84 m diameter wide-chord H-blade (constant chord, rounded tips).

Several one-off projects were undertaken in the US, including the Republic XP-72 (conventional low-wing tractor with Aeroproducts CRP); the Curtiss P-36 (one CRP flight-tested in 1939, model 75E) and the Curtiss-Wright XF14C (1943–1944) with Curtiss propellers. The Hughes XF-11 (1946), with its Hamilton Standard Superhydromatic CRP was unfortunate, as it crashed on its first flight; a second prototype was fitted with single-rotation propellers. The North American XA2J-1 (1949–1952) was a single prototype powered by Allison XT40 turboprops driving two Aeroproducts CRP with large-chord blades ( $c \sim 0.45$ m) and gear-ratio 0.064 [70], shown in Fig. 16. Engine problems were contributing factors in the programme curtailment.

Figure 16. Sketch of the North American XA2J using Aeroproducts CRP.

The Northrop XP-56 (1943) was an aft-mounted pusher CRP with ${N_b} = 3 \times 3$ blades [71], without vertical tail but with a lower safety fin. A first prototype crashed whilst taxiing, and a second one was considered too hazardous to fly, even with design changes; the programme was subsequently stopped. Although the airplane was powered by a P & W R-2800 Double Wasp piston engine, an evolutionary concept was to replace it with a turboshaft with improved thermal efficiency coupled to a counter-rotating propellers [Reference Pavlecka and Northrop72].

The Douglas XTB2D-1 Sky Pirate [73] was remarkable: The aft propeller had a larger diameter than the fore propeller (Hamilton Superhydromatic), something we now understand is to be avoided. The Chance-Vought XF4U-4 was a CRP version (Aeroproducts AD7562-X4) of the very successful F4U-4 (H-S Superhydromatic model 624060), so the advantages of the new configuration were unclear. However, the blades of the aft row were instrumented for vibration testing. The Boeing XF8B was again a one-off attempt by Boeing to enter the carrier-borne aircraft business and designed what turned out to be a one-in-five single-engine aircraft (1944). It used the same Aeroproducts propeller as the Republic XP-72. Only three airplanes were built; by then the US Navy turned its sight to turbojets [Reference Dowgwillo and Svoboda74]. The Curtiss-Wright XP-60 underwent a considerable level of development (-A, -B and -C versions), but ultimately it did not reach production level [75]. Version -C had 3 $ \times $ 3 Curtiss Electric contra-propellers named C(46)615S with 3.71 m diameter.

The Douglas XB-42 (1944-1948) had aft-mounted three-bladed contra-rotating puller propellers (3 $ \times $ 3, 4.12 m diameter; $ \sim $ 900 rpm). The aircraft mounted two side-by-side Allison V-1710 piston engines behind the crew’s cabin, each driving one of the propellers via two long drive shafts [Reference Leonard22] ( $ \sim $ 9 m), with a common gearbox to guarantee operation in case of one engine inoperative, as in a Douglas patent [Reference Burton and Glasgow76]; 1.5 m shafts were hinged at each joint by ball-bearing supports to offload deflections. Due to weight balance considerations, the propellers were placed very close to the tail surfaces, but the aerodynamic interference was believed to be small [Reference Bartlett9]. The aircraft struggled to take-off, due to the low clearance of the vertical fin. Contrary to the claims in said patent, the aircraft experienced severe propeller-induced vibrations, and the programme was cancelled after the second prototype crashed in flight.

There have been a number of other pioneering contributions with a short life. They include Japanese airplanes Kawanishi E15K1 Shiun (1941–1944), with ${N_b} = 2 \times 2$ CRP, reportedly unreliable; and the Kawasaki Ki.64 (1943–1944), with ${N_b} = 3 \times 3$ CRP powered by tandem piston engines, one mounted forward and the other aft of the pilot [Reference Green and Swanborough77]. This solution required a long shaft, as in cases listed earlier.

There were also French airplanes, of course. Alongside the Arsenal VB-10, there was the Sud Ouest SO 8000, with aft-mounted tractor Chauviére CRP, ${N_b} = 3 \times 3$ , and modified pre-war Junkers Jumo engines [Reference Charbonel78]. It allegedly failed to take-off in a few instances, but it eventually flew, with poor handling qualities. It was displayed at the Paris Salon de l’Aviation in 1949.

The North American P-51XR was a 1980s replica of the original 1940s P-51 Mustang; it was powered by a single Rolls-Royce Griffon 57 engine, for air shows exhibitions only.

5.0 Western world: 1950 onwards

The Avro Shackleton was a successful long-range transport and maritime patrol airplane, powered by Rolls-Royce Griffon 57A engines and de Havilland CRP; it served in at least one war theatre, the Suez Crisis of 1956, and it was finally retired in 1991. Three versions of this airplane are known. The propellers were de Havilland $3 \times 3$ constant-rpm, fully feathering, with aft blade row slightly shorter than the fore row. Vibrations appeared around 1,500 and 2,200 rpm, and the airframe was subject to fatigue.

The Fairey Gannet (1959–1979), also known as Fairey 17, was a carrier-borne aircraft of the UK Royal Navy, shown in Fig. 17, widely documented in the technical literature [Reference Turnberg and Brown14, Reference Hollis Williams42, Reference Parzych and Shattuck83]. Its heavily modified AEW Mk.3 had a large radome on the lower fuselage. Both had four-bladed coaxial contra-rotating propellers with Armstrong-Siddeley Double Mamba turboprops and could be started independently.

Figure 17. Sketch of the Fairey Gannet AEW in frontal view. The airplane had a unique Z-folding mechanism for its wings and a large radome fitted under the airframe; it mounted 4 $ \times $ 4 Rotol CRP.

Blade geometry data (chord, twist, thickness, camber for both fore and aft blades) are provided by Ref. [Reference Turnberg and Brown14]. Differences between fore and aft blade design was noted. The aircraft was noted for its unusually high noise, but it was a unique design of interest to Hamilton-Standard because the aircraft could operate each blade row independently [Reference Turnberg and Brown14]. The blade rows had the same diameter (3.76 m), although the aft row was smaller by 3.8cm because the aft barrel had a larger diameter.

The Bristol Type 167 Brabazon was a massive transport airplane of the 1950s that ultimately had no commercial market; it only operated test flights until it was cancelled by the UK government; the aircraft was later scrapped. In bombastic style, the BBC world service broadcast news in several languages of the airplane taking flight, such was the importance of the event in the British Empire. The four contra-rotating propellers were ${N_b} = 3 \times $ 3 with a 4.9 m diameter. The complete details of this CRP are given by the Rotol product brochure [84], and include pitch range and weights.

The engineering concepts explored in the design of the power plant are described in great detail in Refs [Reference Gunston2, Reference Norton85]. Eight Centaurus engines were placed in couples in a V-drive configuration, shown in Fig. 18, although at least six engine layouts ranging from co-axial to parallel and V-drive were considered. They all rotated in the same direction, and contra-rotation was achieved by means of an umbrella gear (aft prop) and spur gear (fore prop). However, the 1940s Bristol Centaurus 18-cylinder air-cooled engines were deemed obsolete for the post-war market, and the project was doomed.

Figure 18. The Bristol Type 167 Brabazon final engine installation; adapted from Ref. [Reference Norton85].

The Saunders-Roe SR-45 Princess (1952), said to be inspired by the Dornier Do-X (a 20-year-old project), was a massive flying boat with six sets of propellers, of which the inner four were CRP and the outer units were single rotation propellers (SRP), shown in Fig. 11.

This arrangement was rather bizarre. A general description of the design is given by Brennan [Reference Brennan86], who only briefly mentions the propellers. This airplane required ten turboprop engines: eight were to be arranged in couples to power four CRP; two were single-engines each powering the outboard SRP. The inboard engines were eight Bristol Proteus 610 turboprops in four pairs. These engines were rated to 3,740 kW when coupled (2 $ \times $ 1,870 kW) plus 7.3 kN of residual thrust [87]. The outboard engines were two Bristol Proteus 600, rated at 1,870 kW $ + $ 3.65 kN of residual thrustFootnote 5. The specification for the Princess was for a 25,000 hp (18,650 kW) power, something that could only be delivered either by large (nonexistent) engines or several of them [Reference Owner88]. Hence the idea of coupling the Proteus engines with CRP.

Laboratory tests for this prototype are recorded in Ref. [Reference Smith, Wright and Owen89], where also off-design issues are reported, such as spray ingestion, propeller stall and damage. Reference [87], which is a report written for the UK Ministry of Supply by the Saunders-Roe company, provides all sorts of specifications, including geometry, aerodynamic data, installed power and propellers. The CRP were manufactured by de Havilland. The gear-ratio was 0.0877:1 for SRP and 0.084:1 for CRP.

Only one aircraft flew (registration G-ALUN, presented at the 1952 Farnborough Air Show), before the programme was cancelled by the UK Ministry of Supply. Many reasons would be invoked, including the fact that corrosion had set in very quickly on stored aircraft. Propaganda films of the time (“Britain’s magnificent Transatlantic Flying Boat”) mention dimensions, power and exceptional comfort. These news reverberated across the specialised press, including a note in Popular Mechanics in September 1949 where it was written that “the plane will have quarters with the comfort and privacy of an ocean liner’s cabin”.

In truth, the airplane did not perform well on flight testing. Only 45 hours of flights were carried out, with two other ships partially built and scrapped; G-ALUN was broken up in 1967. This project showed a serious error of judgement early on in the conceptual design phase; the reality was unkind to the engineers (Sir Arthur Gouge, a “leading authority in this field”) and to the company. Twenty years later, such a ship would have been powered by four turboprops with single rotation propellers, and the passenger arrangements would have been far less luxurious, with the structure being consequently lighter. The age of the seaplane was over by the time the Princess was ready to fly. It is a historical coincidence that in 1952 the first de Havilland Comet, a true innovation, was commencing commercial service.

An important report in 1954 [Reference Gardiner and Brett26] by Gardiner & Brett (chief engineer and chief aerodynamicist, respectively, at de Havilland Propellers Ltd) states that “it has been found that on the payload-range basis the advantage of the counter-rotating propeller is marginal, even for the longest ranges. In view of this and the extreme mechanical complexity of the counter-rotating hub, the current tendency in design is towards single rotation for all types of propellers”.

This view was independently confirmed by Rotol in 1956, when Fairhurst [Reference Fairhurst38], chief engineer, stated that the CRP “has never found favour on civil aircraft, largely on account of its greater expense, complication and more complex service and maintenance problems.” Thus, by the mid 1950s both de Havilland and Rotol were aware that the complexity of the CRP design did not warrant further investment, at least in the United Kingdom. The Fairey Gannet turned out to be the last UK CRP aircraft. It was a successful design indeed. It was manufactured in large numbers, nearly 350 in various versions for several air forces around the world, and served until at least 1979. It was subsequently used for propeller research by Dowty-Rotol and Hamilton Standard [Reference Turnberg and Brown14].

In the US, aeronautical research followed a different path. Among the bizarre concepts tried in those years, the Convair XFY-1 Pogo (1950–1954) is one that deserves a special mention – a new chapter in the history of aviation, according to Convair. A competing design of the same time, the Lockheed XFV-1 Salmon, had the same specifications, same engine and similar Convair CRP. The Pogo was a vertical take-off and landing airplane, named tail-sitter, shown in Fig. 19. It was placed vertically on the ground (sitting on its tail on four wheeled poles) and it would climb thanks to the powerful dual-rotation Convair propellers with ${N_b} = 3 \times $ 3. It was powered by a single Allison YT40 turboprop, which was made of two T-38 side-by-side turboprops connected by a common gearbox [Reference Ginter, Coleman and Long90, Reference Chana91]. The power plant was also unique in that Allison recognised that the tip speeds would be transonic in full flight; hence it proposed to have dual-speed gearbox, with one propeller speed for take-off/hover (gear-ratio 20.4) and the other for level flight/dash speed (gear-ratio equal to 11.9). This change in gear-ratio could only be achieved with idle power [Reference Leonard22].

Figure 19. Sketch of the Convair XFY-1 Pogo experimental tail-sitter and the Northrop XP-56 Black Bullet (different scales).

The Pogo had stability problems, as documented in Ref. [Reference Hollinger and Mitcham92], a NACA report for the US Navy. This airplane was one of the first to be documented in technical reports following wind tunnel and flight tests of scale models [Reference Lovell, Kirby and Smith93]. It achieved at least a couple of remarkable advances: performing flight transition from hover to level flight and back to hover, and flight of a scaled pilotless aircraft. By virtue of the CRP, it did not need another rotor to establish roll and yaw stability in vertical flight. One of the key problems on take-off was the tremendous downwash that would be created by the propellers: It was essentially a helicopter with narrow blades having to propel up to 7,100 kg of weight. In that condition, the downwash would be $ \sim\!40$ m/s, with a mass flow rate through the CRP of $ \sim $ 900 kg/s. By the early 1950s, helicopters had matured to a point where they could provide a variety of aviation services, hence this tail-sitter had no prospect of success. The Pogo performed six transition flights before the programme was halted. Reference [Reference Ginter, Coleman and Long90] has impressive close-up photographs of this singular coaxial VTOL.

However, as a followup, an experimental tilt-rotor V/STOL aircraft was tested. This was the Hiller X-18, shown in Fig. 20, precursor of later tilt-rotors. The aircraft was a hybrid airplane/helicopter with tiltable wings. Both the Pogo and the Salmon required to tilt the complete aircraft. However, it became clear that if propellers were to be used to produce lift and thrust, it would be more rational to maintain aircraft attitude and tilt the rotors [Reference Borst94]: The tilt-wing was born.

Figure 20. Sketch of the Hiller X-18 experimental tilt-rotor with CRP.

The propellers were Curtiss-Electric CRP operating alongside jet vectored thrust. The wing pivoted around 35% chord. This was the first successful attempt at tilt-wing concepts, with engines from surplus Pogo and Salmon projects [Reference Jenkins, Landis and Miller95, Reference Nichols96]. It was powered two Allison YT-40 turboprops and a Westinghouse J34 turbojet that delivered up to 15 kN thrust for pitch control. A number of problems were found, of which we mention the propellers position with respect to the pitch axis, a gyroscopic effect created by wake bending (which increased the power required to tilt the wing), strong ground effects [Reference Newsom and Tosti97] and landing difficulties. It would not be allowed to fly on modern regulations, even experimental type certificates, because it lacked a cross-shaft connection between the engines. Failure on one engine would have been catastrophic.

6.0 Soviet union, 1960s and 1970s

The CRP was particularly successful from the 1960s onwards in the Soviet Union, Russia and Ukraine. The Kuznetzov NK-12 became the engine of choice for quite a variety of airplanes, having been developed over a long time. It is reported that the conception of this engine goes back to the early years after World War II, with Austrian engineer Ferdinand Brandner (1903–1986) playing a major role in the project. The engine was named “NK”, after Nikolay Kuznetzov, chief designer when the product became a reality, although the original name was “TV” and was the result of efforts from post-war German engineers.

First flights with this engine took place in 1954 and 1955. The engine has a 14-stage axial compressor and a 5-stage turbine on a single shaft. The compressors have variable inlet guide vanes to maintain good efficiency across stages and power outputs, and is able to deliver a peak compression ratio $ \sim 13:1$ . This engine series deliver over 11,000 kW of shaft power in its various versions. The NK-12MV is rated to 15,000 hp, or 11,863 kW. Operating at 730 rpm, the corresponding torque would be $ \sim $ 155 kN m.

The Tupolev Design Bureau produced the long-range bomber Tu-95M (1956-present), the Tu-114, the Tu-116, the early-warning aircraft Tu-126 and the Tu-142, all powered by the NK-12 turboprop.

The Tupolev Tu-91 (1955) was a prototype low-wing monoplane, initially intended as a carrier-based bomber aircraft. A single frontal $3 \times 3$ CRP was powered by a centrally mounted Kuznetzov TV-2 turboprop engine via a long drive shaft; the pilots’ seat was between the propeller and the engine (as the Koolhoven F.K.55); the exhausts were behind the wing’s trailing edge.

In the case of the Tu-95, the CRP was ${N_b} = $ 4 $ \times $ 4 (fore and aft, respectively, at 750 rpm), fully reversible props with a 5.6 m diameter, called AB-60. These blades have maximum chord of 0.475 m, a relative thickness of 4.17% and have been designed by TSaGI, the Central Aerodynamics Institute. The weight of propeller and parts, spinners, electrical de-icing system is estimated at $ \sim $ 1,190 kg. The blades are made of aluminium alloy with a central H-spar. The distance between rotor disks is 0.650 m, equivalent to 0.108 diameters. This airplane is still flying in large numbers for reconnaissance and other missions, shown in Fig. 21.

Figure 21. Three Tupolev Tu-95 flying over Moscow in close formation (May 2020).

The Tupolev Tu-114 used similar power plants and CRP [Reference Taylor98]: Kuznetzov NK-12MV and Aerosila AV-60N propellers, shown in Fig. 22. The 5.6 m blades required very long landing gear struts, making the aircraft unusually tall. The nose landing gear strut was over 3 m long. This was a very successful aircraft that operated both USSR domestic and intercontinental flights (including flights to Amsterdam, Tokyo, Montreal) from 1962 until about 1976. There are reports of intense cabin noise at positions near the plane of rotation of the propellers, but no acoustic data are publicly available. The Tu-114 held several aviation records, i.e. FAI record No. 8880 (speed over a closed circuit of 1,000 km with 25,000 kg of payload).

Figure 22. Propellers of the Tupolev Tu-114 (CCCP 76486). Still frame of a video clip at Amsterdam Schipol airport on 1 June 1964. Credit: Nederlands Instituut voor Beeld en Geluid.

The Antonov Design Bureau produced the An-22 (1965 onwards), likewise powered by the NK-12MA turboprops and the same type of propeller blades. The rotors are not synchronised, and the propellers ground clearance is higher than the average ground operators. Figure 23 shows a photograph of the aircraft taxiing on the unpaved runway of Gao airport (ICAO: GAGO), in Mali (2016). Enormous dust ingestion and dust clouds were recorded, but the aircraft could nevertheless deliver large cargo.

Figure 23. Antonov An-22A Antei (Registration UR 09307) landing on dust at Gao airfield, Mali, in December 2016. Still frame of a video clip by the Antonov Airlines.

The CRP architecture was the same as the contemporary Tupolevs, but with a modified propeller AB-90 and a larger 6.2 m diameter, according to Ref. [Reference Kuznetsov, Krymova and Makashov99]. A sketch of the An-22 is shown in Fig. 24 alongside the Tu-95, for comparison. General design aspects of some of these aircraft are described in Ref. [Reference Gordon100], with more detailed documentation in a Russian manual from 1971 [Reference Anikina101].

Following the political crisis in the Soviet Union in the early 1990s, airplanes that were first registered as CCCP- were split between the Russian Federation and Ukraine, and registered as RA- and UR-. The first production airplane was CCCP-56391 (November 1965). The last remaining airplane in service with the Antonov Airlines, registered as UR-09307, Fig. 23, was destroyed in March 2022 during the Russian invasion of Ukraine. A few airworthy An-22 remain in service within the Russian Federation.

Figure 24. Sketch of the Antonov An-22A and Tupolev Tu-95 side by side.

One argument put forward by the design teams is that the availability of large shaft power from the NK-12 turboprop would normally require an increase in propeller diameter. The NK-12 turboshaft appeared to have been developed somewhat independently of the propeller itself. It is likely that the airplanes that used this engine required CRP because this was judged the most reasonable solution to deliver the torque available at the required speed.

The Antonov Bureau also designed the short take-off An-70 (planned in 1984, operational in 1994), powered by four Ivchenko-Progress D-27 three-spool turboprops. This aircraft was intended to modernise the An-12, a four-engined single-propellers aircraft. The Antonov An-70 is the only one, among the examples shown, that has a different blade count between propellers: ${N_b} = $ 8 $ \times $ 6 blades, fore/aft, respectively, at 1,200 rpm, with a 4.5 m diameter, axial separation $x/D \simeq 0.35$ . This CRP is known as SV-27, Figs 4 and 25. The rotor solidity is considerably larger than the previous examples. The propellers have a wide chord with a maximum value of $ \sim 0.60$ m, tapering toward the tips. This leads to an estimated solidity $\sigma \simeq 0.170$ and 0.127 for the fore and aft propellers, respectively.

Figure 25. Sketch of the Antonov An-70 and its SV-27 CRP.

Both power-plant and CRP configuration were used again in the aft-mounted engine of the A-90 Orlyonok, an airplane essentially designed for ground-effect open-water operations. This is sketched in Fig. 26. The airplane is no longer operating. Presumably, flying close to the water caused non trivial risks. A single CRP with ${N_b} = 4 \times 4$ and a 6.2 m diameter was mounted atop the central vertical fin; it was powered by an NK-12MK turboprop, alongside two NK-8-4K jet engines mounted lower, near the bow of the ship. Not much information is available on these propellers, and it is not possible to assess propulsive efficiency and noise.

Figure 26. Sketch of the A-90 Orlyonok central CRP (Aerosila AV-90).

The Ilyushin Il-76LL (1987) was a test bed with a single Lotarev (Ivchenko Progress ZMKB) D-236 engine replacing the inboard right turbofan engine [Reference Yefim and Komissarov102]. This installation is shown in Fig. 27. CRP characteristics included the following: 8 $ \times $ 6 Aerosila blades, 4.20 m diameter, torque split estimated at 41%/59%, rpm variable from 960 (cruise) to 1,100 (take-off/climb-out) and a design 73.4 kN net thrust. This propulsion system was designed to fly at 11,000 m (36,000 feet) at $M = 0.7$ . The blades were made of glass carbon-fibre and carbon composites with a metal spar.

Figure 27. Sketch of the Ilyushin Il-76LL test bed (bottom) with a single prop-fan and Yakovlev Yak-44E AEW with two wing-mounted prop-fans (top).

Likewise, the Yakovlev Yak-42E-LL was a test bed [103, 104] with aft-mounted $8 \times 6$ CRP, with the same engine type, D-236. This prop-fan was mounted on the right side, whilst the left side retained the D-36 turbofan engine in test flights. A further version, the Yak-44E, was an early warning airborne aircraft, of which only one prototype was built. For this version, both engines were prop-fans, with ${N_b} = 8 \times 6$ CRP.

7.0 The 1980s and the fuel crisis

For about 25 years research and development of CRP was almost halted in the western world; the period between 1960 and 1985 saw little activity. In the 1980s, research was scaled up to respond to the high price of fuel at that time. The mechanical complexity of the propellers did not deter the industry from seeking economical solutions at a time of crisis [Reference Strack, Knip, Weisbrich, Godston and Bradley19, Reference Van Zante105].

Modern CRP arising from the mid 1980s onwards are based exclusively on gas turbine engines and are commonly known as prop-fans. The main architectures are geared CRP (both pusher and puller/tractor) and direct-drive CRP. A direct-drive CRP (also known as direct-drive contra-rotating open rotor) only exists in the tractor configuration. This is because the CRP are coupled with the power turbine. A summary of propulsion architectures is displayed in Fig. 28, where there is a mix of geared and ungeared (direct-drive) counter-rotating prop-fans. Configuration (b) is representative of the P&W/Allison 578-DX with the exhaust lobes interfering with the blades. Configuration (d) represents a prototype such as the Kuznetzov NK-110, and configuration (e) is similar to the Kuznetzov NK-108.

Figure 28. Selection of advanced prop-fan configurations.

A summary of data is provided in the supplementary materials. Thanks to the increase in open-source documentation, we have data such as torque split, pitch settings, activity factors, tip Mach numbers and a variety of geometric and aerodynamic data. The propulsion efficiency benefits have been demonstrated in a number of programmes [Reference Harris and Cuthbertson106–Reference Chapman, Godston and Smith108], but they are mostly limited to test beds. These benefits are up to 20% improvement in fuel burn over comparable turbofans. The GE36, the P&W Allison 578-DX, the Ivchenko-Progress ZMKB D-236 and D-27, and the Safran Open Rotor are modern CRP prop-fans.

The P&W/Allison 578-DX prop-fan was tested on a McDonnell-Douglas MD-80 [Reference Howe, Sundt and McKibbon109]. The Allison turboprop was a modified version of a helicopter turboshaft engine (Allison XT701) and used electronic engine controls from P&W. The programme required replacement of one of the JT8D turbofan engines with the 578-DX. CRP characteristics included the following: ${N_b} = 6 \times $ 6 blades rotating at 1,235 rpm, with 3.535 m diameter, axial separation $x/D = 0.186$ , differential planetary gearbox and torque split estimated at $ \sim $ 55%/45%, at flight $M = 0.72$ and altitude of 35,000 feet.

Several problems were encountered both in ground and flight testing, of which we mention a few. To begin with, the exhausts (nine circumferential tubes) were upstream the two rotors; thus, hot gas impingement on the fast-rotating blades was inevitable. This gas impact generated heat loads on the blades and an additional buzzing noise. Furthermore, this demonstrator could not meet One Engine Inoperative (OEI) criteria at take-off if the JT8D-209 engine had to be shut down.

The General Electric GE36 [110] concept was called UDF (Unducted Fan Engine) and was extensively ground-tested in 1985–1986. It demonstrated a fuel consumption rate 20% better than competing turbofans. The engine was then flight-tested on a modified Boeing B727-100 in 1986–1987 [Reference Van Zante105, Reference Harris and Cuthbertson106] jointly by the Boeing Company and General Electric. The baseline Boeing B727 was an aft-mounted tri-jet configuration. For these tests, the GE36 prop-fan engine was mounted on a starboard pylon, whilst the original JT8D engines remained in place at the port-side and the central position. CRP characteristics included the following: 8 $ \times $ 8 blades, 3.557 m diameter, collective pitch range of 86 degrees, 33/29 degree sweep angles (fore/aft rotors), axial separation $x/D = 0.172$ , rotating at 1,390 rpm, with aft blade row operating at $M \gt 1$ at flight $M = 0.72$ and altitude of 35,000 feet. The GE36 engine was 1.25 m shorter than the P&W Allison 578-DX, but 1,140 kg heavier.

NASA’s Advanced Turboprop Programme (1976–1988) deserves a special mention [Reference Hager and Vrabel50], as it provided tests and data on a variety of single-rotation and contra-rotation propellers (Ref. [Reference Hager and Vrabel50] also includes an extensive bibliography). In particular, a single-rotation system named SR-7L consisted of eight highly swept blades, 2.74 m rotor diameter, 41 degree sweep, rotating at 1,700 rpm and generating 40 kN of thrust. It was mounted on the left side of a Gulfstream G-II [Reference Ziemianski and Whitlow111, Reference Graber112] (1986–1988). In fact, considerable work was done to assess single- versus contra-rotation propellers [Reference Anderson113]. En-route noise measurements at a variety of altitudes have been reported in Ref. [Reference Willshire and Garber114], and indicate noise peaks of 70 dB at $M = 0.7$ and altitude of 30,000 feet. This is relatively high in comparison with modern airplanes.

Russian research also moved toward prop-fan and ducted fan concepts. From the late 1980s, Kuznetzov developed the NK-110 [Reference Kuznetsov52]. This engine used a three-shaft architecture with a pusher prop-fan consisting of two co-axial, $4 \times 4$ blades, 4.7 m in diameter, with variable pitch. A three-stage turbine powered the prop-fan through a planetary differential reduction gearbox. The NK-108 variant had tractor propellers. Only papers and drawings are known to exist.

A noteworthy aircraft is the Dornier Seastar CD2 [115] seaplane (1984–2020), which has two propellers in tandem powered by two P & W PT6A turboprop engines ( ${N_b} = 4 \times 4$ , with 2.40/2.35 m diameter).

8.0 Recent developments

A large number of prototypes have been tested in the wind tunnel as part of research projects to provide both aerodynamics and acoustics data, i.e. NASA’s F7/A7 model in Ref. [Reference Woodward116] ( $11 \times 9$ blades, with 0.622/0.607 m fore/aft diameters), the F7/A3 model in Ref. [Reference Woodward117] (same number of blades, smaller aft rotor), and a host of other CRP configurations. NASA’s nomenclature for these systems was to associate a fore blade design number, for example fore F#7, with an aft blade design, i.e. A#7.

One important difference between the designs F7/A7 and F7/A3 is that the latter one has the aft blade row clipped. Other prototypes have been tested in the wind tunnel only for specific reasons, such as blade flutter under high loads, i.e. the F21/A21 CRP ¹¹⁸, ${N_b} = 13 \times 10$ blades, 0.62 m diameter at tip Mach numbers $M = 0.715$ . Several other F#/A# combinations are documented by Ref. [Reference Hughes and Gazzaniga119].

Other configurations tested include the F31/A31, F39/A31, which achieved mixed results. The latter, in particular, had a fore rotor with forward swept blades; it demonstrated poor aero-elastic performance and interaction tone noise levels 8 dB above the F31/A31 reference case [Reference Woodward, Hall, Podboy and Jeracki120]; this was not a good idea. A compendium of this research is given in Ref. [Reference Van Zante105].

More recent designs have been proposed by the CFM consortium and Airbus, such as the A1-PX7 laboratory model [Reference Negulescu121] (11 $ \times $ 9 blades) with a 10% aft-rotor clipping, $x/D = 0.22$ , hub-to-tip diameter ratio equal to 0.35, and design thrust of 10 kN. The data published for this case are vague, as they lack numerical values for context and comparison.

Safran in France has been developing CRP for a number of years, both direct-drive and geared engines. Various levels of multi-disciplinary design and optimisation [Reference Vlastuin, Dejeu, Louet and Talbotec122] have produced extremely complex blade geometries that are swept radially and axially; the new blades look like flags in the wind shown in Fig. 4. Aft-blade clipping is 15% in the latest available version, HERA5. The engine is a derivative of a turbofan engine (Snecma M88) converted into a turboshaft. Unlike the 578-DX prop-fan, the exhaust gas duct is inside the nacelle. Exit is past the aft propeller. The whole propulsion system is mounted on a pylon inclined about 35 degrees from the horizontal plane. This open rotor has been seen spinning for ground testing, but at the present time there is no evidence that it is qualified to fly.

A Rolls-Royce project (company specification RB3011) was a conceptual design having both pusher and tractor configurations ( $8 \times 7$ blades). After wind tunnel aerodynamic and acoustic testing in two different facilities (2008–2010), the project was seemingly left behind (i.e. Rig 145, 1:6 wind tunnel model, build 1 and 2, with varying number of blades, the latter with $11 \times 9$ blades). Acoustic data and installation effects were published in a number of papers.

The contra-vanes concept has been revised in recent research (2015 onwards). This research includes a Safran open rotor (single pusher propeller with variable-pitch swirl recovery vanes, 11 $ \times $ 10 blades), and others, i.e. Ref. [Reference Sinnige, Stokkermans, Ragni, Eitelberg and Veldhuis123] and related papers. In the latter case, a six-bladed rotor with four contra vanes has shown marginal propulsive gains alongside poor aero-acoustic response in laboratory conditions: high tonal components at several multiples of the blade passing frequency have been demonstrated.

9.0 Noise issues

Noise remains one of the unsolved problems of the CRP. The literature on this subject is too vast to be qualified in this context, and it mostly refers to laboratory conditions. Research in the 1950s was modest. After recognising the role of the helical tip Mach numbers, the propeller-fuselage distance, propeller synchro-phasing (mostly for cabin noise annoyance), Ref. [Reference Gardiner and Brett26] acknowledged that “there is little that can be done in the design of the propeller itself to minimise the noise problem.”

On the practical side, there is evidence that the Tupolev Tu-95 is extremely noisy, both outside and inside the cabin, though no measurements prove the actual noise levels; it has been heard from underwater submarines. As reported in Ref. [Reference Kay124], the Tu-95 was “one of the loudest aircraft ever built”, with pilots reporting serious ear discomfort during flight. Yet, it is still flying, as shown in Fig. 21. The Tu-114 was presumably the same, but there are no published data to assess either cabin or external noise. There is evidence that cabin noise was highest on the plane of the rotor, which we would expect. The fundamental blade passing frequency of these propellers is less than 50Hz. Hence noise can be detected by instruments several miles away, with the airplane out of sight.

High noise levels were reported for aircraft such as the Blackburn B-54; again, no measurements are available. Retrieval of air shows bulletins indicates that noise was not a concern back in the 1950s. For example, at the Farnborough Air Show of 1950 both the Bristol Type 167 Brabazon and the Westland Wyvern TF.2 were on display. In the former case, there are comments [Reference Huls125] on the mass balance of the control surfaces, “an offence both against aesthetic and aerodynamic principles” – nothing about noise. The first noise regulations only came into force in 1972 (ICAO Chapter 2). Research has much advanced since then, and the current focus is on prop-fans and open rotors. A sketch of noise propagation and vibration sources is displayed in Fig. 29. Interaction noise between blade rows and between CRP and airframe elements (pylon, fuselage) are recognised as critical contributors to the overall noise and vibration response.

Figure 29. Sources of noise and vibration on a modern aft-mounted prop-fan.

The first acoustic measurements known to this author are those carried out on the Avro Shackleton in the early 1980s [Reference Howell, Bradley, McCormick and Brown126]. Flyover acoustic data were taken, but the actual values are not produced in the research. Interestingly, though, noise from synchronised engines is considerably lower than unsynchronised engines, possibly 10–15 dB, depending on the emission angle. Propeller phase synchronisation has since been an important area of research, i.e. Ref. [Reference Johnston, Donham and Guinn127].

Acoustic measurements were taken by Parzich & Shattuck (1986) on the Fairey Gannet [Reference Parzych and Shattuck83] at full power and 50% power. The aircraft was modified for acoustic testing, and included microphones in the plane of the rear rotor. Data are published on acoustic directivity, front and aft propeller noise split, harmonic content and interaction noise. Flyover noise data indicated showed 98 EPNdB with a microphone placed 1.2 m from the ground. Flight and measurements were carried out following FAR Part 36 regulations. The CRP noise was estimated $ \sim $ 1.5 EPNdB higher than a single propeller with the same output power. A comparison between these flight data and acoustic predictions was shown by Hanson [Reference Hanson128], where a comparison for the first four harmonics is shown, and Parry [Reference Parry129] (without SPL levels).

In a report by the Boeing Company [Reference Shivashankara, Johnson and Cuthbertson130] on the UDF engine it was demonstrated how installation effects can add several dB of tonal noise. This result was confirmed by Airbus and others [Reference Ricouard, Julliard, Omais, Regnier, Parry and Baralo131] on a scaled prop-fan wind tunnel model; at least 5 dB were added by pylon interaction. Russian research on their advanced prop-fans [Reference Kouznetsov, Ganabov, Krymova and Makasho132] reported acoustic loads of 140–150 dB in the aft-empennage, requiring structural upgrades.

It has been now established beyond doubt that clipping the aft blade row leads to a noise reduction [Reference Woodward and Gordon133, Reference Dittmar and Stang134]. From this point of view, configurations with larger aft diameter are well past the state-of-the-art. Likewise, sharp-cut leading edges must be avoided.

The earliest known formulation of CRP noise is attributed to Hubbard in 1948 [Reference Hubbard135]. Acoustics became more focussed in the middle of the 1980s, with a flurry of important research released from that time onwards [Reference Hanson128, Reference Parry129]. This strand of research is now dominating the scientific debate on CRP.

10.0 Critical discussion

The 1950s were characterised by the transition from piston engines to the gas turbine. Even when the turbines proved unreliable (i.e. various Allison T40 versions), progress was relentless and the combustion engine lost its central role. Projects that started with piston engines were converted to turboprops (Northrop XB-35, Blackburn B-54, Douglas XB-42, Westland Wyvern) or cancelled all together (Bristol Brabazon, Saunders-Roe SR-45). Projects that started with CRP (Northrop XB-35, Hughes XF-11), were converted to single-rotation propellers. This process saw the demise of incredible feats of engineering, such as the Napier Nomad turbo-compound, which claimed unparalleled SFC [Reference Chatterton45].

Some failures are attributed to scarcity of time, in spite of laudable resources and creativity. The engineering tools were also limited, and the I-know-best approach prevailed – until it was proved wrong.

A critical flaw in the approach evidenced by both patent and technical literature is that the solution of an individual issue would open the route to a successful development of the CRP. It is entirely possible that a CRP with an advanced turboshaft has a lower SFC relative to a conventional power plant (all other key operational parameters being the same), but the overall direct operating costs are higher. Fuel is not necessarily the highest cost item; quoting fuel prices as a driver for this propulsion system is misleading: excessive maintenance burden may erode any gains from fuel burn savings. The higher acquisition cost of the CRP/prop-fan may well overshadow the cost of a conventional gas turbine engine.

The differences in operating costs, relative to a conventional aircraft, can be estimated from the changes in power-plant operating costs, including fuel burn. Reference [Reference Reynolds136] has a propulsion reliability analysis to establish the likely operating costs and risks for and advanced prop-fan. These estimates are incredibly difficult to carry out, and can only be verified after considerable flight testing. Reference [Reference Reynolds136] also considers the replacement of twin engines under the wing, and the analysis seems to point to nearly the same operating empty weight. After nearly 40 years, this configuration has not materialised.

Today we use conceptual engineering design, an approach based on the rational assessment of top level requirements. At the detailed level, with the multi-disciplinary computational methods available today, it is possible to design an advanced single-rotation propeller that matches or surpasses a CRP on several metrics, including propulsive efficiency, acoustic emissions, overall weight and costs. The latest developments focus on incremental improvement of performance across multiple constraints: acoustics, structural dynamics, aero-thermo propulsion, system integration, operating costs. Some gains have been demonstrated over the baseline of the 1980s designs [Reference Van Zante, Collier, Orton and Khalid137], when such advanced methods did not exist.

The author’s view is that a feasible solution would be a prop-fan that entirely replaces a turbofan on an existing airframe, with minimal changes to the airframe itself. This solution would only be possible for aft-mounted engines (airplanes as the MD-80, which require changes in the supporting nacelles only, unless there are sensible changes in CG position) or as replacement of wing-mounted turboprops (i.e. Bombardier Q400). In fact, parametric analysis of engine placement on the MD-80 indicated that the aft-mounted prop-fans yield the most benefit, including noise. From a safety point of view, prop-fans mounted on aft pylons provide some containment to rotor failure (blade out), because a blade impact with the fuselage would not cause cabin depressurisation [Reference Van Zante105]. For prop-fans mounted under the wings, there would be a need to strengthen the fuselage to contain the blades, alongside additional noise shielding. The limited ground clearance would in any case prevent this solution. Even this is an optimistic assessment, because recent research has indicated that prop-fan and airframe integration is critical to the viability of the CRP. Retrofitting a prop-fan is not a viable option. To the author’s knowledge, no flight testing is forthcoming. Risk aversion has settled in – in sharp contrast with many of the case studies reported.

With regards to the CRP itself, modern designs feature clipped blades on the aft propeller, axial and radial blade sweep to improve load distribution and reduce noise [Reference Majjigi, Uenishi and Gliebe138], increasing axial spacing [Reference Woodward116] and angular speeds, options to trim the CRP to unequal torque. The amount of optimal aft-blade clipping is a function of the aerodynamic load on the fore rotor. Too much blade clipping ends up over-loading the fore rotor and reduces overall propulsive efficiency. In terms of noise reduction, there exist a number of ideas, only tested in laboratory conditions. A summary of noise reduction ideas is provided in Ref. [Reference Dittmar139].

Scanning the technology trends in the 2020s, CRP are increasingly used for a variety of electrically-powered vertical take-off and landing (e-VTOL) vehicles. Unlike gas turbine engines, contra-rotation is achieved by electrical induction, which does not require mechanically complicated gear systems, is more compact, less prone to failure and possibly lighter. However, electromagnetic induction has been proposed for gas turbine engines [Reference Ribarov and Gieras140], and that would appear to overcome some of the mechanical issues with contra rotation. Multiple lightly loaded single rotation propellers are likely to dominate electrical aero-propulsion in the coming years, where fixed-wing and rotary wing aircraft are concerned (NASA X-57 and Volocopter E-Volo, respectively). Highly loaded CRP have been applied to electrically powered VTOL prototypes: Airbus City Bus, EHang 116 and VT-30, Vertical Aerospace VA-X2 and Comac ET120.