We often treat the propeller as a uniform “actuator disc” and the slipstream as a uniform stream tube of accelerated and swirling air. While this simplification is valid for many types of analysis, a propeller is composed of discrete blades. Each blade generates force and also sheds an individual aerodynamic wake downstream.

The discrete nature of the propeller blades has several effects on the propeller, the engine and the airframe.

Blade Loads

The loads on the propeller blades are not constant in flight. There are several conditions that change the aerodynamic environment of a blade as a function of its angular position along the orbit of the propeller. This variation produces a cyclic change in the force generated on the blade.

Angle of Attack

As we saw in last month’s Wind Tunnel about P-factor, the lift on the blade can vary as it goes around its orbit, even if the propeller is rotating at a constant rpm. There can be several causes for this variation in blade lift.

The first is the effect of the angle of attack of the airplane on the angle of attack and airspeed the blades encounter over a full rotation of the propeller.

A positive airplane angle of attack tilts the prop axis relative to the wind. Individual propeller blades encounter different conditions at different points in the propeller rotation. An up-going blade has a lower blade AOA and airspeed than if the shaft were parallel to the wind. A down-going blade has a higher blade AOA and airspeed. This causes the lift of each blade to vary cyclically within a period of one cycle per revolution.

Wake Impingement

If the wake of any part of the airframe impinges on the propeller, it affects the forces on the blades. This is particularly true on pusher configurations because the propeller is some combination of the wing, tail or engine support structure.

The frequency of the perturbations to the forces on a blade depends on how the blade passes behind an airframe component. For example, if the engine is pylon-mounted, as is the case on the Lake Amphibian, each blade passes through the wake of the pylon once per revolution. If the engine is in line with a wing, like on the Rutan canard designs such as the Long-EZ or on an airplane with wing-mounted pusher engines like the Piaggio Avanti or Beech Starship, each blade passes through the wing wake twice per revolution.

Upwash and Downwash

A lifting wing deflects the airstream. There is an upwash ahead of the wing and a downwash behind it.

On a multi-engined airplane with wing-mounted engines and tractor propellers, the props are immersed in the upwash ahead of the wing. This means that the propeller axis of rotation is tilted upward relative to the local airflow at the plane of the prop, even if the propeller shafts are physically aligned with the direction of flight. The blades will encounter the same variation of local angle of attack and airspeed around the prop orbit as we saw in our discussion of the effect of airplane angle of attack.

The downwash behind a lifting wing changes the angle of attack of the parts of the airplane behind the wing. On a pusher configuration, wing downwash affects the inclination of the oncoming airstream relative to the propeller axis of rotation. This produces the same type of once-per-revolution variation of blade forces as we saw as a result of airplane angle of attack.

Effects of Blade Force Variations

The aerodynamic effects we have just discussed produce cyclic variations of the loads on each propeller blade. The characteristic frequencies of the variations in blade load are multiples of the propeller rate of rotation.

Resonance and Blade Fatigue

The cyclic changes in blade thrust can excite flapping oscillations of the blade. This can be a significant problem if the natural frequency of a flapping mode of the blade is at or near the same frequency as the excitation caused by the fluctuation of the aerodynamic loads on the blade.

If the two frequencies couple, the blade will experience a resonance. The cyclic aero load will excite the structural flapping mode of the blade, causing the flexing of the blade to grow and the transient forces on the blade to increase far above the initial magnitude of the aerodynamic load. This can cause fatigue failure of the blades and is a particular problem for metal propellers or props with metallic shanks holding on wooden or composite blades.

It’s important that the propeller blades be stiff enough so that their natural frequencies are well separated (usually higher frequency) from at least one-per-revolution and two-per-revolution blade-passing frequencies. Certified propellers are tested extensively to ensure that they do not suffer from resonance issues within their normal operating range. There have been problems in the past with Experimental props with metal blades that were not so extensively tested, and at least two early attempts at variable-pitch props for homebuilts had to be withdrawn from the market after a series of in-flight blade failures.

Cyclic Loadings From the Prop

In addition to affecting the blades themselves, the cyclic variations of blade loading are transmitted into the propeller hub, the prop shaft and eventually the airframe via the engine and engine mount. This creates structure-borne vibration in the airframe and also cyclic loads that can fatigue mounting bolts and other structural components.

The frequency of these vibrations is different than that experienced by a single blade because it also depends on the number of blades. In essence, the characteristic frequency is multiplied by the number of blades. If, for example, a single blade experiences a one-per-revolution variation of load, a two-blade prop will put a two-per-revolution force variation into its hub and a three-blade prop will generate a three-per-revolution excitation.

The number of blades also affects the amplitude of the vibration caused by blade-force variations.

The amplitude of the propeller-induced vibration is a function of the loading of the individual blades. The cyclic loading on each blade is proportional to the nominal or average load the blade is carrying.

To the first order, the thrust generated by the propeller is divided equally between the blades. Increasing the number of blades reduces the load carried by each individual blade and therefore reduces the amplitude of the cyclic variation of load on each blade.

The result of these phenomena is that as the blade count increases, the frequency of the propeller-induced vibrations increases and its amplitude decreases. With more blades, the propeller will generate higher-frequency, lower-amplitude vibration.

The lower amplitude and higher frequency of the prop-induced vibrations are one reason to use a prop with more blades. In general, using fewer blades and more diameter is more efficient until the blade tip Mach number becomes the limiting factor on diameter, but the reduction in vibration amplitude with more blades is often worth the slight reduction in efficiency. Also, higher-frequency vibrations are less objectionable to people, so both effects make the propeller seem “smoother” to the occupants of the airplane.

The post Prop Blade Effects appeared first on KITPLANES.

Torque

The engine exerts torque on a shaft to turn the prop. There is an equal and opposite torque reacted through the engine mounts into the airframe. This torque tends to rotate the airplane in the direction opposite to the rotation of the propeller.

Torque reaction is a major consideration in the design of helicopters, where the shaft is oriented vertically to drive the rotor. Most helicopters use a tail rotor that pushes sideways on the end of the tail boom to counteract the torque driving the main rotor. Without the tail rotor, the body of the helicopter would spin opposite to the rotation of the main rotor.

The torque of the engine produces a pure rolling moment on the airplane, unlike P-factor, which produces yawing moments. Aerodynamic rolling moments generated by the wings and ailerons oppose the rolling moment due to engine torque.

Torque is most significant at high power and low airspeed. The torque of the engine, and hence the rolling moment it exerts on the airplane, is at its highest at maximum throttle. The ability of the airframe to generate aerodynamic force is proportional to airspeed squared, so the ailerons have the least power to fight engine torque at low airspeed.

Two flight conditions where this is the case are right at liftoff, with the engine at full power to start the climb, and at the beginning of a missed approach, when the pilot applies full power to stop the descent and start climbing again.

For most general aviation airplanes, the rolling moment due to engine torque is small enough relative to the control power of the ailerons that it is easy to control. If the airplane has a single engine and is very powerful, the picture changes. Some WW-II fighters had so much torque that the ailerons could not overcome the torque if the pilot abruptly applied full power at low airspeed. The F4U Corsair was particularly susceptible to this because it was a Navy fighter that could fly slowly enough to land on the aircraft carriers of the period, but also had a very powerful engine to give it the top speed and high rate of climb it needed to be an effective fighter and interceptor. If the pilot slammed the throttle forward suddenly to initiate a wave-off, the airplane would roll uncontrollably. This caused more than one accident in service, and the problem persisted when civilians who were used to lower-powered airplanes started flying surplus fighters after WW-II.

The problem of torque causing difficulties with roll control reappeared when people started to install turbine engines on some composite kit airplanes originally designed around piston engines. Turbines with 400 to 500 hp were installed, replacing 160- to 200-hp piston engines. There were several accidents with these early turbine conversions where the airplane lifted off at full power and the pilot was unable to keep the airplane from rolling with the engine torque after liftoff.

In these high-powered singles, the correct technique is to realize that they might have a minimum control speed (Vmc) below which its only safe to use partial power. The airplane can still fly and climb at airspeed below this because the engine has so much excess power that even at part throttle it provides enough power to climb, but it is unsafe to use full power until the airspeed is above the critical speed.

It’s important to note that the fact that torque is so significant on high-powered airplanes does not mean that P-factor is not. P-factor rises with increasing power also, so in addition to the pure rolling moment caused by torque, the pilot must still contend with the yawing moment due to P-factor.

More About P-factor

One component of P-factor is the yawing moment induced by the propeller. The slipstream spirals around the fuselage and generates an angle of attack on the vertical fin. This causes the fin to generate side force that pushes on the tail, causing a yawing moment. This phenomenon occurs on single-engine airplanes and is a direct function of power and airspeed. The yawing moment comes from the effect of the slipstream on the airframe, not directly from the propeller itself.

There is a second component of P-factor that can cause the propeller to directly exert a yawing moment on the airframe. This component is a function of both power and angle of attack. It’s caused by how the thrust of a propeller blade varies as a function of where in the rotation it is.

At zero angle of attack the propeller shaft is parallel to the airstream. A propeller blade that is in the part of the rotation where it is moving upward sees the same blade angle of attack and airspeed as a blade that is in the part of the rotation where it is moving downward.

Placing the airplane at a positive angle of attack tilts the prop axis relative to the wind. Now the blades see different conditions as they move around the propeller orbit. An up-going blade sees a lower blade AoA and airspeed than if the shaft were parallel to the wind. Conversely, a down-going blade sees higher blade AoA and airspeed.

Because of this asymmetry in blade angle of attack and airspeed, an up-going blade generates less thrust than a down-going blade. This moves the centroid of the thrust of the propeller laterally from the axis of rotation toward the down-going blade. If the prop is on the airplane’s centerline, this thrust offset generates yaw in the direction of prop rotation, toward the down-going blade. This effect gets more pronounced as angle of attack increases, so it will be at its maximum when the airplane is flying at minimum airspeed and maximum power.

This component of P-factor is responsible for the “critical engine” phenomenon on twin-engine airplanes. If an engine rotates so the down-going blade is outboard of the shaft, this shifts the thrust outboard. If an engine rotates so the up-going blade is outboard of the shaft, this shifts the thrust inboard. If both engines rotate the same way, the one with the down-going blade outboard will make more yaw because its thrust is further outboard of the center of gravity. This makes the other engine, with the down-going blade inboard, the “critical engine.” If it fails, the opposite-side engine with the down-going blade outboard makes more yaw because of thrust offset. The airplane will have a higher Vmc if the “critical engine” fails than if the other engine fails.

The cure for this is to arrange for the propellers of the two engines to rotate in opposite directions (counter-rotate) so that the down-going blade is inboard on both sides.

High-powered multi-engine airplanes like WW-II bombers, twin-engine fighters and piston- or turbine-engine airliners typically have propellers driven by reduction gearboxes. Different gear sets were used to drive left-hand-rotating and right-hand-rotating propellers so the airplane could have counter-rotating propellers.

Many general aviation twin-engine airplanes have direct-drive engines with no gearbox. As light twins appeared, they typically had engines that rotated the same way, giving them a critical engine. This was enough of a concern that Lycoming eventually developed and offered direct-drive engines with reverse rotation. Once these engines appeared, an airplane could have counter-rotating props without gearboxes by using a left-hand-rotation engine on one wing and a right-hand-rotation engine on the other. The Piper Seminole and Twin Comanche CR are examples of this.

The post Torque and P-factor appeared first on KITPLANES.

The propeller of an airplane generates thrust that drives it through the air. In addition to propelling the airplane, the thrust of the propeller affects both stability and trim. The relationship between the thrust line and the center of gravity can produce throttle-dependent effects on the trim of the airplane and also affects its stability.

Thrust Line

The thrust line is a line along which the thrust of the propeller acts. To the first order, the thrust line is coaxial with the axis of rotation of the prop and normal to the plane of the prop. Both the vertical position of the thrust line and the inclination of the thrust line relative to the centerline of the airplane affect how thrust will affect the stability, trim and control of the airplane.

Thrust-Induced Moment

When a force acts on a body, it will produce a moment about the center of gravity unless it is acting directly through the CG. The moment will be proportional to the thrust and the distance between the thrust line and the CG measured normal to the thrust line.

Equilibrium

When the airplane is in steady-state trimmed flight, the moments generated by the thrust are balanced out by aerodynamic moments generated by the flying surfaces of the airframe. If the thrust line does not pass directly through the CG, then the horizontal tail must generate an offsetting pitching moment to trim the airplane.

Throttle-Dependent Pitch Characteristics

When the pilot moves the throttle and changes the thrust of the propeller, it disturbs the equilibrium of the airplane. The thrust changes immediately when the throttle moves. Two things happen: Thrust no longer equals drag, so the airplane begins to accelerate or decelerate, and the thrust-induced pitching moment changes in response to the change in thrust.

Consider the case when the pilot increases power: The increase in thrust increases the moment being generated by the thrust of the propeller right away. At the instant of the thrust increase, the airplane begins to accelerate, but the airspeed has not yet increased in response to the thrust increase.

The aerodynamic moments counterbalancing the moment generated by the thrust do not change until the airspeed does unless the pilot moves the stick. This means that during the transient period when the airplane is accelerating, the moments will be out of balance and there will be a net moment in the direction of the thrust-induced moment.

If the thrust line passes above the center of gravity, then the initial throttle-induced transient will be nose down, and if the thrust line passes below the center of gravity, the initial transient will be nose up.

A reduction in thrust will produce the opposite effect: If the thrust line passes above the center of gravity then the initial throttle-reduction induced transient will be nose up, and if the thrust line passes below the center of gravity, the initial transient will be nose down.

This does not take into account the effect of the propeller slipstream flowing over the airframe. For example, the effect of the propeller slipstream over the tail can produce powerful throttle-induced effects that can either compensate for or amplify the direct effect of the thrust change.

Speed Stability

Power is equal to thrust times velocity. This means that at constant power, the thrust of a propeller is inversely proportional to airspeed. The faster the airplane goes the lower the thrust, and the slower the airplane goes the higher the thrust.

This change in thrust alters the thrust-induced moments. The faster the airplane is flying, the smaller the thrust-induced moments are. The slower the airplane flies, the bigger the thrust-induced moments are.

Aerodynamic moments also vary with airspeed. They are proportional to airspeed squared, so they increase as the airplane goes faster and decrease as the airplane slows. What this means is that starting from a trimmed equilibrium, if airspeed increases, the thrust-induced moment decreases and the aerodynamic moment increases. If airspeed decreases, thrust-induced moment increases and aerodynamic moment decreases.

This phenomenon affects what is called the speed stability of the airplane. If the thrust line passes above the center of gravity so that the thrust-induced moment is nose down, then the airframe must generate a nose-up aerodynamic moment to trim it.

In this situation, the thrust effect is stabilizing. As the airplane goes faster, the thrust-induced nose-down moment gets smaller and the aerodynamic nose-up moment gets bigger. This causes the airplane to nose up in response to an increase in airspeed.

It will also nose down in response to a decrease in airspeed as the thrust-induced nose-down moment gets larger and the aerodynamic nose-up moment gets smaller. This is a stable condition since the airplane will pitch in a direction that opposes the initial change in airspeed.

The opposite is true if the thrust line passes below the center of gravity. In this situation, the net thrust-induced moment is nose up and it must be trimmed by nose-down aerodynamic moment. As the airplane goes faster and the thrust decreases, the nose-up moment generated by the thrust also decreases while the nose-down aerodynamic moment increases. This will cause the aircraft to nose down and increase speed further.

A decrease in airspeed will cause the thrust-induced nose-up pitching moment to increase and the aerodynamic nose-down pitching moment to decrease, causing the airplane to nose up and decelerate. This is an unstable condition since the pitch response will be in a direction that amplifies the initial perturbation in airspeed. This speed stability effect is additional to the fundamental pitch stability of the aircraft, which is determined by how the aerodynamic moments change in response to a change in angle of attack.

Thrust-Induced Moment Issues

In general, excessively high thrust lines are more likely to cause problems than excessively low thrust lines. This is in part because there are useful configurations (like amphibians and pusher ultralights) that use high-mounted engines to keep the propeller clear of the surface or airplane components.

High Thrust Line

On an airplane with the thrust line significantly above the CG, high thrust and low airspeed tend to drive the nose down since it is difficult for the tail to generate enough moment trim to counter the thrust-induced moment. This can cause difficulties on takeoff and lead to a very dangerous hop-off phenomenon. If the tail cannot overcome the thrust-induced moment and rotates the airplane nose up at an airspeed below liftoff speed, then the airplane can get going fast on the runway at full-up or nearly full-up elevator. When the speed gets high enough for the tail to rotate the nose up, the rotation can be very abrupt and lead to the airplane lifting off suddenly with a high nose-up pitch rate. This phenomenon has caused several accidents in early flight testing of new designs with high thrust lines.

Up and away, a high thrust line adds to speed stability, but it can also lead to a dangerous condition if the engine loses power during initial climb. Because the thrust-induced moment is nose down, an engine failure will cause a sudden nose-up change in pitching moment along with the loss of thrust. This means the pilot will find themself in a situation where the nose is going up and the airspeed is already decaying due to the loss of thrust. This greatly increases the danger of a stall and loss of control of the airplane. This phenomenon was one contributor to the very poor safety record of the BD-5. The BD-5 had a high thrust line, an unreliable engine and sensitive controls and was very light so it had relatively little inertia to help maintain speed when thrust was lost. Several BD-5 accidents resulted from an engine failure on initial climb followed by a pitch-up and a stall.

This phenomenon can also be an issue for ultralight pushers with a high thrust line since they are lightweight and high drag so that airspeed will decay very rapidly in the event of an engine failure. A nose-up pitch transient caused by the loss of the thrust of a high-mounted pusher engine will exacerbate this unless the pilot pushes the nose down aggressively in response to a loss of power.

Low Thrust Line

While problems induced by low thrust lines are rare, there is one notable recent example. The Boeing 737 MAX initially exhibited a thrust-induced pitch-up when approaching a stall in the flaps-up configuration. Boeing responded to this by incorporating an automatic trim system to compensate for the pitch-up. This system had a failure mode that caused two serious accidents and led to the grounding of the MAX until the system was redesigned to eliminate this failure mode and recertified.

The post Thrust Line Effects appeared first on KITPLANES.

The propeller slipstream affects the aerodynamics of the airplane. A propeller accelerates air backward and also imparts a swirl. The air in the slipstream spirals aft, rotating in the same direction as the rotation of the propeller. Both the higher airspeed in the slipstream and the swirl affect the forces on the portions of the airframe that are immersed in the slipstream.

P-Factor

On a single-engine airplane with a tractor propeller, the fuselage and tail are fully immersed in the propeller slipstream. This means that on an airplane with a conventional (by American standards) propeller rotation (right blade moving down, left blade moving up), the air in the slipstream is moving left to right on top of the fuselage and right to left on the bottom of the fuselage.

The vertical fin is on top of the fuselage, so it is immersed in the left-to-right flowing portion of the slipstream. The sidewash from the cross flow of the slipstream induces an angle of attack on the vertical tail that causes it to develop lift (side force) to the right. This side force causes a nose-left yawing moment. This phenomenon is called P-factor.

The intensity of the P-factor is a function of airspeed and power setting. At high power and low airspeed, as is typical of climb, the P-factor is most powerful because the magnitude of the swirl in the slipstream is greater relative to the airspeed of the airplane. This maximizes the sidewash angle on the fin, and thus its angle of attack and lift coefficient. The effect is increased by the fact that the slipstream velocity is at its highest relative to the airspeed of the airplane so that the side force on the fin is at its maximum relative to the other forces on the airplane.

At higher airspeed and lower power, as is typical of cruise, P-factor is significantly lower than in climb. Power off, like in glide or approach, P-factor is negligible. The variation of P-factor with flight condition means that the rudder deflection needed to keep sideslip at zero and maintain coordinated flight varies with airspeed and power setting.

It is possible to set up the airplane to maintain coordinated flight with zero rudder pedal input and force at a single flight condition by using either a fixed rudder trim tab or by mounting the fin at a slight yaw angle to counteract the P-factor. At every other flight condition, the pilot must deflect the rudder with pressure on the rudder pedals to keep the ball centered unless the airplane has a rudder trim system the pilot can adjust in flight.

Absent adjustable rudder trim, the designer must choose what flight condition to set up the airplane for zero rudder force to trim. In most cases this will be cruise since the airplane will be cruising most of the time, and the pilot should not have to hold any continuous control pressure to maintain coordinated cruising flight.

Most airplanes need right rudder deflection and pressure on the right rudder pedal to keep the ball centered in climb because the P-factor is greater at high power and low airspeed. If the airplane is set up for zero rudder pressure coordinated cruise flight, it will need more rudder deflection to keep the ball centered in climb.

AoA Differential

The spiraling slipstream behind a propeller produces an upwash behind the ascending blade and a downwash behind the descending blade. If there is a wing behind the prop, this effect increases wing AoA behind the ascending blade and decreases it behind the descending blade. At angles of attack below the stall, this increases lift on one side of the prop and decreases it on the other. This lift differential causes a rolling moment that can help counteract the torque of the engine.

At higher angles of attack, the asymmetry in local AoA caused by the slipstream swirl can cause an asymmetric stall where the wing downstream of the upgoing blade stalls before the wing downstream of the descending blade. On a single-engine airplane this asymmetric stall will happen near the wing root and will likely be relatively benign as long as the outer portions of the wing do not stall first.

On a twin with wing-mounted engines the effect can be more dramatic. More of the wing is immersed in prop wash because there are two engines, and the slipstreams of the propellers are farther outboard than on a single. If both engines rotate in the same direction, there will be a significant prop-wash-induced difference in the angles of attack of the outer panels of the two wings. The farther outboard the AoA difference is, the more likely it is to induce an asymmetric stall that leads to a roll-off into an incipient spin.

Slipstream Convection

If an airplane yaws relative to the oncoming airflow, the slipstream convects downstream and is turned by the oncoming flow. This moves the slipstream laterally on the airframe. On a single-engine airplane the slipstream will move toward the retreating wing, increasing the lift of the retreating wing, and off of the advancing wing, decreasing the lift of the advancing wing.

On a twin, the slipstream will move outboard on the retreating wing and inboard on the advancing wing.

On a single, the movement of the slipstream increases the lift of the retreating wing and reduces the lift of the advancing wing. On a twin, the movement of the slipstreams shifts the lift of the retreating wing outboard and the lift of the advancing wing inboard. Both effects will cause a rolling moment toward the advancing wing, rolling the advancing wing down. This is an unstable variation of rolling moment with sideslip. If this unstable moment is not compensated for by a stable moment generated by the rest of the configuration, the airplane will have an unacceptable unstable dihedral effect.

The effect of slipstream convection is greater on a twin than on a single because more of the wing is washed by slipstream. Also, on a single, the fuselage acts as a fence that impedes turning of the slipstream, while on a twin the slipstream is freer to move laterally with sideslip of the airplane.

This phenomenon caused a serious issue with the last series of civilian transports built by the Glen L. Martin company. The Martin 2-0-2 was a twin-engine transport designed to be a replacement for the war-surplus DC-3s (C-47s) in widespread use at the time. In flight test, the 2-0-2 proved to be laterally unstable at high power and low airspeed, particularly with the flaps down. This presented the Martin designers with a serious problem. The cure for lateral instability is to increase wing dihedral. Since the problem was discovered in flight test, after the tooling to produce the wing carry-through structure was already complete, there was no economically viable way to increase the dihedral of the whole wing.

Martin’s solution was to introduce a dihedral break in the wing where the outer panels attached to the engine nacelles. This cured the lateral stability problem, but the hastily engineered wing attach system had structural issues that led to a 2-0-2 shedding a wing in flight. The airplane was fixed and improved and reintroduced as the Martin 4-0-4. The 4-0-4 was modestly successful, but is notable primarily as the last civilian airplane built by Martin.

The post Propeller Effects–Lateral Directional appeared first on KITPLANES.

In preliminary design and for performance calculations, we treat the engine, propeller and airframe as separate entities. The only engine effect we typically consider in preliminary performance analyses is cooling drag. We neglect the propeller’s effect on the drag of the airframe and treat the prop as a pure thrust producer. Any effect the propeller might have on drag is kept as a small reduction in propeller efficiency.

For jet aircraft, the question of what is thrust and what is drag can get quite complex since the flow though the engines affects the airflow on the airframe significantly, particularly for aircraft with buried engines like fighters and some other military jets. Thrust-drag accounting is a major effort in and of itself on these airplanes and is far beyond the scope of what we need to consider for simple propeller-driven airplanes.

That said, although we can treat the performance effects of the propeller airflow very simply in most cases, the effect of the prop wash on the aerodynamics of the airplane in flight is not trivial. In a recent message, a colleague commented that “Installing a prop is like attaching a tornado to the airframe.”

Slipstream

In order to produce thrust, a propeller accelerates a column of air backward. The air in this column downstream of the prop is called the slipstream. In the slipstream, the air is moving faster than the free stream, and it is also rotating in the direction of the rotation of the propeller. The streamlines in the slipstream are spiral in form as the air flows downstream.

Both the added velocity and the spiraling swirl in the slipstream can affect the aerodynamic behavior of the airplane. The slower the airplane is flying, the greater both the slipstream velocity and swirl for a given horsepower to the propeller. Accordingly, slipstream effects are at their highest at low airspeed and high power.

Stall Characteristics

On most tractor airplanes, some portion of the wing is immersed in the propeller slipstream. For a single, the root of the wing is in the prop wash, while for a twin it is the portions of the wing behind the engines. The slipstream has two effects, both of which tend to delay the stall. First, the propeller turns the oncoming air and directs it aft more parallel to the prop shaft than the free stream. This reduces the angle of attack of the wing behind the prop. Second, the increased velocity in the slipstream increases the lift the immersed portion of the wing is generating.

Because it delays the stall of the blown portion of the wing, the slipstream changes where the wing stalls power-on. Particularly for a single-engine airplane, the point of initial stall on the wing will shift outboard of the propeller slipstream. A wing that had a benign root-first stall power-off may well stall farther outboard and tend to roll off during a power-on stall.

The combination of the delay in the stall and the shift in stall characteristics is what makes departure stalls more abrupt and dynamic than power-off stalls. The airplane stalls at a slower airspeed and at a more nose-up attitude, making the stall break more dramatic than with power off, and the lateral shift of the initial stall increases the chances the airplane will depart laterally into an incipient spin.

Lift

The propeller-induced velocity in the slipstream increases the airspeed of the portion of the wing and flaps it blows over. The lift generated by the wing and flaps is proportional to airspeed squared, so at low airspeed and high power the lift increase due to the additional slipstream velocity can be quite large. For STOL takeoffs, the combination of high power and powerful flaps can generate a lot of lift and reduce liftoff speed and takeoff roll significantly.

There are two concerns with using this additional propeller-induced lift on takeoff. First is that while the flaps generate a lot of extra lift in the prop wash, they also generate a lot of drag. The airplane must have enough excess thrust to climb after lifting off on prop wash-augmented lift.

The second concern is that if the prop wash goes away, so does the excess lift. This can lead to a very unsafe condition shortly after takeoff. The airplane is dependent on the extra lift from the propeller slipstream to keep flying. An abrupt throttle cut or engine failure will leave the airplane stalled in a nose-high attitude at an airspeed significantly below its power-off stall speed. It will require a significant loss of altitude to accelerate to a high enough airspeed to establish stable un-stalled flight.

Pitching Moment

As the slipstream flows aft, it flows first over the wing and then over the tail of the airplane.

As it flows over the wing and (sometimes) flaps, the slipstream is deflected downward. As we have just seen, this generates some additional lift, but it also generates some nose-down pitching moment.

Downstream of wing and flap, the down-deflected slipstream impinges on the tail. The tail sees both higher airspeed and a more negative angle of attack as a result. These two effects induce an additional down load on the tail and cause a nose-up moment as a result. The slipstream-induced nose-up moment of the tail is usually larger than the nose-down moment induced on the wing and flaps, so adding power at low airspeed will tend to cause the airplane to pitch nose up.

On airplanes with high-mounted engines like amphibians and some pusher ultralights, this slipstream-induced down load on the tail can help compensate for the nose-down effect of the high thrust line. If the tail is placed directly aft of the propeller so it is blown by the prop wash, then the slipstream-induced down load on the tail will generate a nose-up moment that helps compensate for the nose-down moment produced by the high thrust line.

Tail Power for Takeoff Rotation

Prop wash over the tail significantly increases elevator power at low airspeed. This helps with lifting the nosewheel off the runway to rotate for takeoff. It helps the pilot make a smooth takeoff if the elevators can lift the nose gear at well below the stall speed. The pilot can initiate a gentle rotation, starting somewhat before liftoff speed, and get a smooth, progressive transition from rolling to flying.

T-tail configurations that get the tail out of the slipstream sometimes have poor takeoff characteristics because the elevators can’t rotate the airplane until it’s going faster. Full up is often required to rotate. Once rotation starts, the pilot will have to react by reducing elevator deflection to prevent over-rotation or hop-off. Several T-tailed certified airplanes suffered from this problem, and at least one type started out with a conventional low-mounted horizontal tail, was converted to a T-tail in a later (supposedly improved) model and then converted back to a low-mounted tail again in a still later revision. The final low-tail version of the airplane is still in production.

Next month we will turn our attention to the lateral/direction effects of the propeller slipstream.

The post Design Process: Slipstream Effects appeared first on KITPLANES.

The engine of an airplane must produce enough power to meet the requirements of all phases of the mission.

First, the engine must accelerate the airplane up to liftoff speed during the takeoff roll.

Once the airplane is airborne, the engine must provide enough thrust to maintain flight and enough additional thrust beyond that to make the airplane climb.

In cruise or loiter, the engine must generate enough thrust to overcome the drag of the airplane and maintain level flight.

Ideally, the power needed to provide adequate performance for all three of these segments of the flight should dictate approximately the same size engine, but this is not always the case. The requirements for each phase of the flight are different, and which one drives the size of the engine will vary depending on the mission and specific configuration of the airplane.

Takeoff

During the takeoff roll, the engine must accelerate the mass of the airplane from rest to liftoff speed. The length of the takeoff roll is determined by the liftoff speed and the acceleration of the airplane.

Liftoff speed is determined by the stall speed of the airplane. The higher the liftoff speed, the longer the takeoff roll.

Acceleration during the takeoff roll is determined by the power-to-weight ratio. This ratio compares the maximum power of the engine to the weight of the airplane. The lower the power-to-weight ratio, the slower the acceleration, and the longer the takeoff roll.

While the aerodynamic drag of the airplane does reduce acceleration somewhat near the end of the takeoff roll, drag is less significant during the takeoff roll than during climb or cruise.

Relatively small changes in either liftoff speed or acceleration can have a surprisingly large effect on how long the takeoff roll is. This is because the distance required to accelerate to a given speed varies quadratically with both speed and acceleration.

For a given acceleration, the distance required to achieve a given speed is proportional to the target speed squared. For example, doubling liftoff speed would increase takeoff roll by a factor of four.

Similarly, the distance required to reach a given speed is inversely proportional to acceleration squared. Because of the squared relationship between distance and acceleration, cutting the acceleration in half will multiply the distance by a factor of four.

We see the effects of these phenomena when we compare ultralights or STOL aircraft to other types. The reason STOL aircraft use so little runway to take off is that they have low liftoff speeds and high power-to-weight ratios. They accelerate rapidly to a relatively low liftoff speed. With both squared functions working in their favor, they take off in a very short distance.

Climb

Once the airplane is off the ground and climbing the picture changes. A significant portion of the power developed by the engine now goes to generating thrust to offset the drag of the airplane and maintain airspeed. It is only the available power above that required to maintain level flight, otherwise known as excess power, that goes to driving the airplane uphill to make it climb.

The rate of climb of the airplane is determined by a quantity called “specific excess power” (Ps), which is the excess power above that required to maintain level flight divided by the weight of the airplane.

For the purpose of engine sizing, we can start out assuming constant weight and then look at the effect of changing power on rate of climb.

The excess power available is considerably less than the total power of the engine. For example, on a typical light trainer, the excess power at best rate of climb will typically be between one-fourth and one-third of the rated power of the engine.

This makes rate of climb much more sensitive to total engine power than other aspects of the airplane’s performance. Changing the rated power of the engine does not change the power required to maintain level flight, so any change in the power delivered by the engine goes directly into excess power. The percentage change in rate of climb will be much larger than the percentage change in rated power. For example, for the typical light trainer, increasing rated power by 25% will double the excess power available and therefore double the rate of climb (100% increase in climb).

Cruise

Power required to maintain airspeed is a direct function of the drag of the airplane. The higher the L/D ratio of the airplane, the lower the power required to cruise at a given speed.

To the first order, the power required in level flight is proportional to the cube of the airspeed. The first thing to realize from this is that simply adding power is not a very effective way to make an airplane go much faster. Doubling the power of an airplane will only make it go about 25% faster.

It is far more effective to use aerodynamic drag reduction to make an airplane go faster than it is to simply increase the power.

If we look at these two factors, we can get an idea of how cruise considerations will size the engine.

High-drag airplanes will never go very fast and may require quite a bit of power to achieve an acceptable cruise speed.

Airplanes designed to fly fast will require a relatively high power-to-weight ratio, even if they have low-drag airframes, because of the V-cubed relationship between speed and power.

Matching the Engine

The challenge for the designer is to take all of these considerations into account. Sometimes the three considerations—takeoff, climb and cruise—all dictate approximately the same size engine.

Light Sport Aircraft and trainers are a good example of this. They do not fly very fast and cruise speed is of secondary importance, so once the engine has enough power to give them acceptable takeoff and climb performance, it is sufficient for cruise. In other cases, however, one consideration dominates.

STOL airplanes are dominated by takeoff roll and initial rate of climb. They usually have relatively high drag configurations that will not fly very fast no matter what power is available. Accordingly, they tend to be overpowered for a reasonable cruise. They need a high power-to-weight ratio for takeoff acceleration and lots of excess power for initial climb in spite of their high drag, but they will cruise at a relatively low throttle setting for fuel economy.

Fast airplanes will often have engines sized by their top-speed requirements. Even with an aerodynamically clean, efficient airframe, high speed requires high power.

This can prove advantageous for takeoff and climb. High-speed airplanes tend to have high wing loadings and high stall speeds that would tend to increase takeoff roll. The extra horsepower needed to achieve the high cruise speed the airplane is designed for will also improve takeoff acceleration and help keep the takeoff roll down to an acceptable distance.

There is likely to be a large mismatch in power required for the three flight conditions in the case of an aerodynamically efficient but relatively slow airplane like a motorglider. Because of their low drag, these airplanes can successfully maintain level flight with very low power for their weight. An extreme example of this is the English Electric Wren, which is able to carry a pilot using an engine rated at 3.5 hp.

If the engine is sized to match this low-power cruise condition, the airplane is likely to be quite deficient in power for the takeoff roll, where the aerodynamic efficiency of the airplane does not matter and acceleration is strictly determined by the power-to-weight ratio.

Likewise in climb, where excess power over weight determines performance, an engine sized for cruise will not produce much excess power. Even at full throttle, the rate of climb is likely to be sluggish.

The trade here will really depend on the mission of the airplane. A self-launching sailplane will have an engine sized for takeoff and climb. The primary purpose of the engine is to replace the tow plane, not to cruise efficiently. The airplane is going to spend most of its time with the engine shut off and stowed or the prop feathered.

If an airplane is intended to cruise slowly or loiter efficiently under power, then the engine will be sized small to produce minimally acceptable takeoff and climb performance.

The post Design Process: Engine Size appeared first on KITPLANES.

On some airplanes, the basic empennage does not provide acceptable flying qualities in all flight regimes. Supplemental surfaces to augment directional stability, improve high angle of attack characteristics or improve spin characteristics can solve these issues.

One class of such surfaces is typically mounted low on the aft fuselage.

Ventral Fins

Ventral fins are vertical fins that extend downward from the bottom of the fuselage. They are typically smaller, with a lower aspect ratio, than top-side vertical fins, primarily because they need to be short to maintain acceptable ground clearance.

A big advantage of ventral surfaces is that they are in clean airflow, particularly at higher angles of attack. This means they retain their effectiveness in situations where the other tail surfaces may be immersed in the wake of other parts of the airplane.

Relatively small ventral fins can increase directional stability more than their area alone would suggest for two reasons:

First, the ventral fin increases the effective span of the whole vertical tail, so in addition to generating side force itself, the effective span increase due to the ventral fin also increases the effectiveness of the whole vertical tail.

Second, the ventral fin acts as a fence that increases the side force that the fuselage side above the ventral fin generates in a sideslip.

Ventral fins are a common feature of floatplanes. The vertical fins on these airplanes were sized for the original landplane configuration. Floats are directionally destabilizing because placing them so that their center of buoyancy on the water is at the airplane’s center of gravity puts them far enough forward that their lateral aerodynamic center is forward of the CG. This effect may be large enough that an increase in fin effectiveness is necessary to maintain good flying qualities. One way to add the needed extra directional stability is by adding a ventral fin on the centerline of the bottom of the fuselage below the original fin and rudder.

In addition to improving directional stability, a ventral fin improves yaw damping and spin recovery. Ventral fins are particularly effective at improving spin characteristics because they are in clean airflow at high angles of attack. They provide yaw damping that slows the rotation rate in the spin and makes it easier for the rudder to break the spin.

One example of this is the evolution of the tail of the Cessna Skycatcher LSA. The first prototype was lost in an unrecoverable spin during testing. (The pilot bailed out successfully and was not injured.) The second version of the airplane still had problems recovering from spins. The final modification to the tail was the addition of a small ventral fin. This cured the spin recovery problem.

Afterbody Strakes and Canted Ventral Fins

Horizontal Strakes

Low-aspect-ratio horizontal strakes mounted to the aft fuselage are often used to improve longitudinal stability at high angles of attack. The strakes act as an additional low-aspect-ratio horizontal tail. Because of their low aspect ratio, the stabilizing effect at low angles of attack is relatively small, but it increases significantly at higher angles of attack.

If the goal is to increase pitch stability in cruise flight or move the aft CG limit of the airplane aft, strakes are relatively ineffective. A better solution is to increase the span and area of the horizontal tail.

Where strakes are most effective is at higher angles of attack. This makes them useful for improving stall recovery and stability near the stall.

Afterbody strakes produce a stabilizing nose-down pitching moment. The moment increases as angle of attack increases up to and beyond the angle of attack at which the wing stalls. This nose-down moment helps limit the angle of attack of the airplane and aids in getting the nose down to break an incipient stall.

Low-aspect-ratio surfaces have nonlinear lifting characteristics. Not only does their lift increase with increasing angles of attack, but the slope of the lift curve also increases as angle of attack increases. They achieve their maximum lift at significantly higher angles of attack than higher-aspect-ratio surfaces.

At low angles of attack, the flow over a strake is fully attached and its lift varies linearly with angle of attack.

At higher angles of attack, the flow separates from the leading edge and outer edge of the strake and forms a vortex. The vortex generates lift differently than attached flow. As angle of attack increases, the strength of the vortex increases. The lift it generates is proportional to both the angle of attack and the strength of the vortex. Because of this, the lift of the strakes varies with AoA squared in the vortex-lift AoA range.

The nonlinear lift of the strakes means that the stabilizing effect of the strakes gets larger as angle of attack increases into the vortex-lift regime. This makes them ideal devices to compensate for the pitch-up and high AOA loss of longitudinal stability that is sometimes a problem on T-tail and V-tail configurations. The nose-down moment of the strake replaces the lost nose-down moment from the tail and helps the airplane maintain pitch stability as it approaches the stall.

One example of this application of strakes is on the Cessna Citation Mustang, which has a T-tail and a pair of afterbody strakes mounted low on the aft end of the fuselage.

Canted Dual Strakes

Some airplanes use a pair of aft body strakes mounted with a significant anhedral angle to form an inverted “V.” The anhedral angle moves the strakes down relative to the rest of the airplane. This keeps them clear of the exhaust of fuselage-mounted engines and helps ensure that they will be in clean airflow, out of the wake of the wing or forward fuselage. It also allows the strakes to do double duty. They improve directional stability and also have the high AOA pitch stabilizing effect of horizontally mounted strakes.

An example of the use of canted dual strakes in combination with a T-tail is the Learjet 60.

The Cirrus Vision Jet has a pair of strakes below the fuselage. These strakes have anhedral that mirrors the dihedral of the V-tail surfaces so the combination of V-tail and strakes form an X-tail.

Afterbody Strakes for Drag Reduction

Another application of dual strakes is to control the flow on the fuselage afterbody to reduce the drag of the airplane. Dual strakes are common on airplanes that have upswept afterbodies like those of military transports. If the aft portion of the fuselage is swept upward, either for an aft door or to give more ground clearance from a tail strike on the runway, the flow can form a pair of vortices that shed aft from the fuselage. Adding a pair of properly oriented strakes to the afterbody can reduce the strength of these vortices and control their position so they shed cleanly from the aft fuselage and produce less drag.

The post Design Process: Strakes and Ventral Fins appeared first on KITPLANES.

The majority of airplanes use the conventional tail configuration with a single, centrally mounted vertical tail and rudder. Sometimes, however, the vertical tail is separated into two or more surfaces.

One such configuration has vertical fins mounted to the tips of the horizontal tail. These fins can be the only vertical tails on the airplane or can be used in combination with a center vertical.

Directional Stability Improvement

Tip fins on the horizontal tail are sometimes added to an airplane with a single vertical tail to increase directional stability. Often this is a result of a modification to the airplane that makes the original vertical fin inadequate to properly stabilize the airplane. Tip fins are common features of floatplanes. The floats themselves are directionally destabilizing, so adding floats may make an increase in fin area necessary to maintain good flying qualities. One way to add the needed extra fin is by putting tip fins on the horizontal tail.

Hangar Height

On an airplane with multiple vertical fins, each fin is smaller than the single fin the airplane would otherwise need. Splitting the fin into two or three parts reduces the overall height of the tops of the fins and hence the overall height of the airplane.

Over time, as airplanes got bigger the height of their fins increased accordingly. At the same time, in the WW-II era, tricycle landing gear began to replace tail-dragging gear for larger airplanes. This moved the tail of the airplane up when at rest and further increased the overall height of the plane at rest. The result was that a configuration with a single vertical tail would not fit into existing hangars. Designers incorporated multiple fins on such airplanes as the B-24, B-25 and Lockheed Constellation to keep the tip of the fins low enough to fit into the hangars available at the time.

Structural Considerations

A multi-fin tail has both structural advantages and disadvantages. The overall weight of the tail system (horizontal + vertical) will be a function of the relative importance of these for the specific configuration.

On the positive side, each tail is smaller, so the bending moment on the fin spars is less. This advantage is increased with a full H-tail where the fins are mounted to the horizontal tail near the vertical centroid of the fin so that the length of cantilever for the fin spar is only half of what it would be for a similar fin mounted at its root.

An additional structural advantage is that the centroid of the fin area is much lower, so the aerodynamic side force generated by the fin produces lower torsional loads on the fuselage than a single central fin cantilevered upward from the top of the fuselage would.

The primary structural disadvantage is that the horizontal tail must have additional structure to attach the tip fins and also additional structure to carry the aerodynamic loads on the fins.

Aerodynamic Effects

Pitch

The fins mounted on the tips of the horizontal tail act as end plates. They increase the lift-curve slope of the horizontal tail, which increases both the stabilizing effect of the horizontal tail and the control power of the elevators. This effect may allow the horizontal tail to be smaller for the same level of stabilizing effect.

Directional Stability

The fins of a multi-fin tail are typically lower aspect ratio than the single vertical tail that they substitute for. The fins at the tips of the horizontal tail also are not end-plated by the fuselage like a single centrally mounted fin is. Both of these effects reduce the lift-curve slope of the fins. This means that for a given total fin area, the fins of the multi-fin tail produce less directional stabilization than a single higher-aspect-ratio fin. Accordingly, the total fin area needed for the multi-fin tail will be larger than for a single vertical configuration.

Wake Effects

For aircraft with bluff bodies that shed turbulent wakes, a central vertical fin may be immersed in the wake of the upstream components and be unable to stabilize the airplane. Mounting the vertical fins at the tips of the horizontal tail moves the fins out of the aerodynamic wake and places them into clean airflow where they can function effectively to stabilize the airplane.

This situation is common on pusher aircraft, which tend to have bluff afterbodies upstream of the tail. On gyroplanes, for example, it’s common to have a three-fin configuration. The central fin has the rudder and is immersed in the propeller slipstream so the rudder is effective. The tip-mounted verticals on the horizontal tail provide most of the directional stabilization.

Gyroplanes also tend toward multi-fin configurations for geometric reasons. The acceptable height of the vertical fin (s) is limited because the fins are below the rotor and the fins must be short enough so that when the rotor blades flap as they spin they remain clear of the fins.

Spin Characteristics

The H-tail configuration with the fins mounted on the horizontal tail generally has good spin recovery characteristics. The portion of the fins below the horizontal tail remain in clean air at high angles of attack and retain their effectiveness. Also, the advancing fin in a spin is moving into clean air as the airplane rotates and provides effective anti-spin yaw damping.

Slipstream Effects

Moving the vertical fins off of the centerline of the airplane changes the way the fins interact with the propeller slipstream. This can be used to either move the fins out of the prop wash or deliberately immerse them in it, depending on the configuration of the airplane and the designer’s intent.

On a single-engine tractor airplane, the swirl in the slipstream caused by propeller rotation is a primary cause of P-factor on an airplane with a conventional tail. Because the vertical fin is above the fuselage but not below, the fin sees some sidewash from the swirl in the slipstream. This generates side force on the fin, causing a yawing moment that the pilot must compensate for with rudder. This effect is most pronounced at the high power settings and low airspeeds typical of climb.

Mounting the fins on the tips of the horizontal tail takes them out of the propeller slipstream and eliminates this effect. This is why the Ercoupe has an H-tail configuration. Eliminating most of the P-factor was key to making the airplane compatible with two-control operation.

On a multi-engine airplane, using an H-tail deliberately places the fins and rudders in the slipstream of the propellers, which can increase the rudders’ control power. In an engine-out situation, having a rudder in the slipstream of the operating engine increases the rudder power available to offset the yawing moment caused by the loss of thrust and drag of the failed engine.

Mechanical Considerations

The control linkages for a tail with the fins and rudders mounted outboard on the horizontal tail are more complex than for a single vertical fin/rudder configuration because the controls must be routed outboard through the horizontal tail to the rudders.

Care must also be taken in the layout of the elevators and rudder to ensure that the rudders do not clash with the elevators when a rudder deflects inboard. There are several ways to accomplish this. The rudder hinge line can be aft of the elevator trailing edge, or the tips of the elevators can be cut away at an angle to clear the deflected rudders. Another approach is to arrange the rudders so they deflect outboard only. This works well mechanically but sacrifices some rudder control power.

A three-fin configuration with fixed outboard fins and the rudder on the central fin eliminates these issues and can work well, provided the single central rudder has enough control power. This arrangement of a central (often all-moving) rudder in the prop wash combined with outboard fixed fins to augment directional stability is a common feature of modern gyroplanes.

The post H-tails and Triple Tails appeared first on KITPLANES.

A T-tail is a configuration where the horizontal tail is mounted on top of the vertical tail rather than directly to the fuselage. T-tails are frequently used on military transports and also on sailplanes.

Potential Advantages

The primary aerodynamic reason to mount the horizontal tail on top of the fin is to move it out of the wake of the wing. Ideally, the horizontal tail will see less downwash from the wing and may also see a bit higher dynamic pressure since it is clear of any reduction in energy caused by the wing’s passage through the air.

In theory both of these effects could allow the horizontal tail to be smaller and reduce the aerodynamic interference between the wing and the tail.

Another aerodynamic effect of the T-tail configuration is that the horizontal tail can act as an end plate for the vertical tail and increase the lift curve slope of the fin. This can allow the vertical fin to produce more directional stability for the same area and may allow the fin to be slightly smaller to provide the same directional stability as a conventional fin.

If the vertical tail is swept back, the sweep will move the tip of the fin, and hence the horizontal tail, aft relative to the aft end of the fuselage itself. This increases the tail arm and thus the stabilizing effect of the horizontal tail.

Disadvantages

A T-tail is more complex both structurally and mechanically than a conventional tail. The pitch control linkages will be more complex since it is necessary to run the controls up inside the vertical fin to get to the horizontal tail and actuate the elevator.

There are also multiple structural issues. The first is that the vertical tail must withstand the aerodynamic loads generated by the horizontal tail and elevator in addition to the lateral force generated by the fin itself. The end-plating effect of the horizontal tail also shifts the centroid of the side force the fin develops upward, which increases the bending moment on the fin.

A second structural concern is aeroelastic. The horizontal tail has a large yaw inertia. Mounting it on top of the fin means that the yaw inertia of the horizontal tail dominates the torsional modes of the fin. This drives down the natural torsional frequency of the tail and makes it more flutter-prone. The fin must have higher torsional stiffness to prevent flutter.

The combination of the added loads on the vertical fin and the need for much higher torsional stiffness means that the structure of the T-tail will be significantly heavier than the structure of a conventional tail.

Aerodynamic Concerns

Single-Engine Nosewheel Liftoff

On a single-engine tractor-propeller airplane, the T-tail configuration moves the tail and elevators above the slipstream of the propeller in addition to moving them out of the wing wake. This can be a problem on takeoff since the elevators don’t have the benefit of prop slipstream to help them lift the nosewheel off the runway to rotate for takeoff. This has proven to be an issue on several production airplanes with T-tails that ended up with longer takeoff distances because the airplane had to be rolling faster before the pilot could rotate the nose up for takeoff.

Junction

From a structural design viewpoint, it would be ideal to have the main spar of the vertical fin intersect the main spar of the horizontal tail directly. This ideal structural design produces an aerodynamic interference problem: The spars of the surfaces are normally located at or near the maximum thickness point on the chord. The problem is that an intersection between two airfoils with the maximum thicknesses in the same position is likely to experience flow separation aft of the maximum thickness points as the skin of both airfoils pulls away from the airflow simultaneously.

This problem can be reduced by staggering the maximum thickness points of the fin and horizontal tail fore and aft, but to do this requires additional structure to connect the spars of the two surfaces together.

Many T-tailed airplanes have some sort of junction body or fairing at the intersection between the fin and horizontal tail to smooth the airflow over the junction and fair any additional structure needed to connect the fin and the horizontal tail.

High AOA

T-tail configurations sometimes suffer from significant nonlinearity in pitch characteristics with changes in angle of attack.

At low angles of attack, the wing downwash is relatively low and the horizontal tail is above the majority of the downwash. As angle of attack increases, two things happen. The downwash angle in the wake behind the wing increases, and the horizontal tail moves down relative to the oncoming flow, into the downwash behind the wing.

The increased downwash produces a negative angle of attack increment on the horizontal tail. This negative angle of attack change on the tail produces a nose-up pitching moment increment. The result of this is that the airplane loses pitch stability as angle of attack increases.

How severe this effect is depends greatly on the details of the airplane’s geometry. Some T-tailed jet airliner designs have had severe issues with this phenomenon. The British BAC Trident had a fatal accident during flight testing when the pitch-up caused by the T-tail placed the airplane into an unrecoverable deep stall. The horizontal tail on the original DC-9 was enlarged significantly from its initial size before the airplane first flew because Douglas designers were concerned that the DC-9 might be vulnerable to the same sort of problem.

In Practice

In general, T-tails have not proven to give any meaningful advantage for typical single-engine light airplanes. There are some successful T-tailed light airplanes, but there is no strong evidence that they have better performance than similar airplanes with conventional tails.

Several production light planes with T-tails had issues with pitch nonlinearity and nosewheel liftoff. One manufacturer, Piper, actually went full circle during the evolution of its product. The original PA-32 Cherokee Six had a conventional low-mounted all-moving horizontal tail. The next iteration, the PA-32RT-300 Lance II, had a T-tail. The Lance then evolved into the PA-32-301 Saratoga, which reverted to the low-mounted tail of the original Cherokee Six.

T-tails are used successfully on several military jet transports. The C-141, C-5 and C-17 all have high wings and powerful flaps. They use T-tails to keep the horizontal tail clear of the wakes of the flaps and engines.

T-tails are also nearly universal on modern high-performance sailplanes. The primary reason for this is more operational than aerodynamic. Sailplanes have a single main wheel landing gear and a tail skid or tailwheel. At rest, one wingtip rests on the ground. The aft fuselage is small in diameter, so a low-mounted horizontal tail would either be very close to the ground or actually touch the ground with the glider in the wing-down resting position. Using a T-tail moves the horizontal tail up enough so that it is safely clear of the ground.

The post Design Process: T-Tails appeared first on KITPLANES.

A V-tail is a configuration where the horizontal stabilizer and vertical fin are replaced by a pair of surfaces mounted at a high dihedral angle (usually about 45°). Because these angled surfaces can produce both vertical force and side force, the two V-tail surfaces can replace the three surfaces of a conventional tail. The movable surfaces on a V-tail (called ruddervators) move collectively for pitch control and anti-symmetrically for yaw control.

V-Tail Concepts

At first look, it appears that by getting two surfaces to do the work of three, the airplane can have a smaller total tail area, which should reduce both drag and weight. This idea is still cited as a reason to use a V-tail configuration. In fact, with one small exception, this isn’t true in reality.

Tail Area: V-tails generate side force when the airplane is sideslipping and pitch force in response to an angle of attack change, thus generating stabilizing moments in pitch and yaw. Likewise, deflecting the ruddervators symmetrically produces a pitching moment, and deflecting them anti-symmetrically produces a yawing moment to control the airplane.

Two common misconceptions are that the effective area of the V-tail in yaw is the side-projected area of the tails, and the effective area of the V-tail in pitch is the top-view-projected area of the tails. These ideas are the foundation of the belief that the total tail area could be smaller with a V-tail than a conventional tail.

Unfortunately, this is not the case because the aerodynamic force generated on a V-tail surface acts normal to the plane of the surface, not directly sideways for yaw or vertically for pitch. Accordingly, only part of the force the tail surface generates actually stabilizes or controls the airplane, while the rest is canceled out by an equal and opposite force in that axis generated by the opposite-side V-tail surface. For example, in pitch the V-tails generate symmetric forces normal to the tails. Part of this force is oriented vertically and contributes to pitch control or pitch stability. The other component of force is oriented horizontally, and those forces are equal and opposite on the two tails and provide no net contribution to pitch stability or control.

Similarly, the yaw lateral component of force on each tail does contribute to yaw stability and control, but the vertical components are equal and opposite and generate a rolling moment that may or may not be desirable.

The net result is that in order to provide the same stabilizing influence in pitch and yaw as a conventional tail, the total area of the V-tail surfaces must be the same as the total area of a properly sized vertical fin and a properly sized horizontal tail. This means there is no net reduction in tail area or wetted area of the airplane using a V-tail.

The wasted force (e.g., the lateral component when trying to trim in pitch) means that the V-tail surfaces need to generate more total aerodynamic force to produce a given trimming or stabilizing influence. This results in extra induced drag, which increases trim drag and also the loads on the tail structure.

There is no reduction in skin friction drag because there is no reduction in wetted area, and there may be some increase in trim drag because of the cost of the “wasted” aerodynamic force on the tails.

There is a weak argument that because a V-tail has fewer junctions between the tail surfaces and the fuselage there will be some reduction in junction and interference drag.

Structure: The V-tail has significant structural disadvantages with respect to a conventional tail. Because of the dihedral angle between the tails, there is no continuous carry-through structure like there is on the horizontal stabilizer of a conventional tail. The V-tail surfaces individually are larger than either the vertical tail or the horizontal tail panels of a conventional wing tail. Both of these factors make the structure of a V-tail heavier.

As we have just seen, the aerodynamic forces on a V-tail will be larger than those on a conventional tail because only part of the total load on the V-tail panels actually produces stabilizing and controlling influence on the airplane. The wasted force we just saw still adds load that the tail structure must withstand, even if it does not help stabilize or control the airplane.

Flutter: V-tails are more prone to flutter than conventional tails, as sadly illustrated by some of the problems experienced by early model Beech Bonanzas. Because the left and right ruddervators are not interconnected and can deflect anti-symmetrically, they can produce large torsional forces on the fuselage. This makes a V-tail airplane more susceptible to a flutter mode where fuselage torsion couples with anti-symmetric flapping of the ruddervators. It’s particularly important to ensure that the control surfaces of a V-tail are properly mass balanced.

Overall, the V-tail is likely to be heavier than a conventional tail and will probably have the same or higher drag than a conventional tail producing the same total stabilization and control power.

Stability and Control Issues

Adverse Roll: When the ruddervators are deflected anti-symmetrically to command yaw, the aerodynamic forces on the tails are up and sideways on one tail and down and sideways on the other tail. The sideways forces act in the same direction and produce the commanded yaw. The vertical forces produce an uncommanded roll in the opposite direction to the commanded yaw. This large rolling moment due to yaw-control deflection of the tails also produces large torsional loads on the fuselage and is likely to force the fuselage structure to be heavier.

Dutch Roll: V-tails produce a lot of yaw/roll coupling due to their high dihedral. This can reduce Dutch roll damping. The usual cure for low Dutch roll damping is to increase directional stability. If we try to do this by enlarging a V-tail, we also increase the yaw/roll coupling, which works against the desired effect of enlarging the V-tail. It might be necessary to adjust the dihedral of the wing to help offset the rolling moment produced by the V-tail.

Pitch-Up: As the angle of attack of the airplane increases, the V-tail surfaces move down relative to the wing. This can immerse them in the downwash of the wing, which will produce a negative angle of attack change on the tails. This produces download on the tail, which drives the nose up. V-tail configurations can be prone to nonlinear pitch behavior and pitch-up at the stall.

Why Use a V-tail?

A V-tailed airplane will probably be heavier than an airplane with a conventional tail. V-tails have no drag advantage, are more flutter prone and introduce the possibility of pitch stability issues. In spite of this, there are some special cases where it can be a good design choice.

Tail Ground Clearance: Some sailplanes have V-tails to give the tail adequate clearance from the ground when the glider is on the ground with one wingtip resting on the ground.

Wake of Upstream Components: Some airplanes have features that produce a turbulent or separated air wake in the area where a conventional horizontal tail would normally be mounted. A V-tail is one approach that can move the stabilizing surfaces up and out of this wake. While this can work, there is always the risk that the V-tail will move down into the wake at high angles of attack.

Fewer Parts: A V-tail has two panels instead of three, and they can sometimes be made symmetrical left/right, which simplifies the manufacture of the airplane. On an airplane with a powered control system, a V-tail configuration can have fewer (albeit larger) actuators than a conventional tail.

A final reason to use a V-tail has nothing to do with the aerodynamics of the airplane. Several military airplanes (notably the Northrop Tacit Blue and YF-23) have V-tails to reduce their radar signature.

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