ISO 27001

The CG is considered the balancing point of a body for weight and balance purposes, as will be discussed shortly. The CG is determined by summing moments about a datum and dividing by the weight. When the CG is not aligned with the mechanical axis, the cyclic control must be sufficiently displaced to compensate the unbalanced CG condition. The helicopter fuselage will be tilted so that the heaviest end or side will be lower in a hover. Subsequently changing the CG of the aircraft will require the cyclic control to be repositioned. If cargo, fuel, or personnel are loaded or unloaded, the new CG will require compensating with cyclic. An aft CG will require forward cyclic and forward CG will require aft cyclic. Corresponding lateral cyclic inputs are required for lateral CG displacements. The limit of cyclic authority plays the most important role in determining the CG limits of a helicopter. However, the CG limit is not defined by the full displacement of the cyclic; the limit must be maintained within the cyclic authority to ensure adequate control and a margin of safety.

If the safe CG limits are exceeded, the aircraft will enter uncontrollable flight. Full cyclic displacement will be unable to compensate for the extreme CG, and the aircraft will roll or pitch in the direction of the extreme CG, likely resulting in aircraft damage or destruction

(uncommanded right roll, low G flight/mast bumping, LTE).

CG Limits. Stability and controllability are greatly affected by an aircraft’s CG. The CG is defined as the imaginary point at which the total weight of a body can be considered to be concentrated. In fact, if a line was attached to an aircraft at its CG, the airframe could be lifted without rotating because it would be balanced evenly about that point. In a similar manner, the CG of an aircraft will determine the fuselage attitude of the aircraft in both a hover, and forward flight, and will determine the amount of controllability and stability of the helicopter about all three axes.

To calculate the location of the CG in the design phase, engineers determine the weight of each component and its location relative to the centerline, waterline and nose of the aircraft. Then, all of the components' moments relative to the CG are determined by multiplying the weight of the component by the distance to a datum point of the aircraft to get a total moment. The total weight of all the components is added up to get a total weight. We can then divide the total moment, in units of inch-pounds (in.-lbs.), by the total weight, in units of pounds (lbs.), to get the distance of the CG from the aircraft’s datum point. In the case of the TH-57, the datum is

defined as the nose of the helicopter, and the moment arms are measured in inches behind the nose of the aircraft. A moment is determined by multiplying the moment arm (inches) by the weight in that particular area (passengers, fuel, baggage, etc.). Once the moments are summed, the sum is divided by the total weight, and this quotient will be the arm of the CG behind the nose in inches. When all is said and done the designers ensure that the CG is in its most optimal position for the missions and flight profiles of the aircraft. In general, a more forward CG, within limits, results in a more stable aircraft, since the main rotor's contribution to AOA instability is decreased.

By the military specification against which aircraft are tested, a minimum of ten percent (sometimes 12-15 percent) control margin must exist for flight at a CG limit. The CG limits, then, are established such that the pilot has at least the minimum control margin available when flying at a CG limit. What must be recognized is that the helicopter CG limits are established for a static, stabilized flight condition, where dynamic maneuvers are not involved. In order for the pilot to stop or reverse a dynamic maneuver when operating at a CG limit, it will be necessary to use some or all, of the available remaining control margin. If a pilot is operating at or near a CG limit it is imperative to maintain a very "stable" flight profile by limiting the amount of aggressive or dynamic maneuvers.

CG limits are affected by the type of rotor system, (Figure 7-8). A semi-rigid rotor that has a small CG range can be improved by using a longer rotor mast. A fully articulated or rigid rotor has its CG range established based upon control authority through the hinge point or effective hinge point. Tandem rotor helicopters are very versatile cargo aircraft because their CG range, originating between the two rotors, is very broad in the cabin area.

Figure 7-8 Effect of Rotor Systems on CG Travel 708. HELICOPTER STATIC STABILITY AND MANEUVERABILITY

The helicopter is inherently unstable because it tends not to stay in equilibrium. Since the helicopter is unstable, the pilot has to spend much of his time keeping the helicopter on the desired flight path. Many helicopters are now built with auto flight control systems so that, with greatly reduced pilot workload, they will maintain heading, trim, and/or altitude in forward flight. However, the TH-57 does not have such a system and therefore requires constant vigilance in both the hover and in-flight regimes.

Equilibrium occurs when the sum of the forces and moments around the CG are equal to zero.

An aircraft in equilibrium will travel in a constant direction at a constant speed, developing no moments that would cause it to rotate around the CG. Since an aircraft can rotate around three different axes, we must consider its stability around each of these axes. Lateral stability is stability of the lateral axis around the longitudinal axis (roll). Longitudinal stability is stability of the longitudinal axis around the lateral axis (pitch). Directional stability is stability of the longitudinal axis around the vertical axis (yaw). Each motion requires a separate discussion.

Some basic assumptions must be made to simplify this discussion. First, the disturbances will be small enough to keep the change in pitch attitude, and degree of yaw and roll small enough so that the aircraft does not approach any unusual attitude. Disturbances are external and not caused by the pilot. The pilot applies no inputs to correct the displacement from equilibrium.

Any moment that corrects the airplane's attitude is the result of the design of the aircraft.

Any discussion of aircraft stability requires an explanation of how the rotor system, fuselage, vertical stabilizer, horizontal stabilizer, etc, affect the longitudinal, lateral, and directional stability of the airplane. This is critical to understanding why the rotor disc is tilted or coned a particular way, why the tail is where it is, and why the vertical stabilizer is as big as it is. This discussion will revolve around conventional helicopters, that is, helicopters with a single main rotor and tail rotor.

A helicopter’s maneuverability is the ease with which it will move out of its equilibrium position. Obviously, maneuverability and stability are opposites. A stable aircraft tends to stay in equilibrium and is difficult for the pilot to move out of equilibrium. The more

maneuverable an aircraft is, the easier it departs from equilibrium, and the less likely it is to return to equilibrium. If an aircraft needs to move quickly from its trimmed equilibrium attitude, it will have weak stability. Of course, this means the aircraft will be more difficult to fly in equilibrium and will require more of the pilot's attention. The mission of a specific aircraft dictates the compromises between stability and maneuverability the designer will have to make.

With a basic understanding of static stability, it is possible to discuss each aircraft component and its individual contribution to static stability. Afterwards, the effect of all the components can be used to discuss the overall static stability of the aircraft.