Aircraft Weight and Balance Handbook

Arm ............................................................................2-2 Moment ......................................................................2-2 The Law of the Lever .....................................................2-2 Shifting the Balance Point or CG...................................2-4 Solution by Chart .......................................................2-4 Basic Weight and Balance Equation ..........................2-4 Solution by Formula ...................................................2-5 Mean Aerodynamic Chord .............................................2-5 FAA-Furnished Weight and Balance Information ........2-6 CG Range ...................................................................2-6 Empty Weight CG Range (EWCG) ...........................2-6 Maximum Weights .....................................................2-6 Number of Seats .........................................................2-6 Maximum Baggage ...................................................2-6 Fuel Capacity..............................................................2-6 Oil Capacity (Wet Sump) ...........................................2-6 Data Pertinent to all Models .....................................2-10 Manufacturer-Furnished Information ..........................2-10 Chapter 3 Weighing the Aircraft and Determining the Empty Weight Center of Gravity ........................3-1 Introduction ....................................................................3-1 Requirements .................................................................3-2 Equipment for Weighing................................................3-2 Preparation for Weighing...............................................3-3 Scale Preparation ........................................................3-3 Weigh Clean Aircraft Inside Hangar ..........................3-4 Equipment List ...........................................................3-4 Ballast .........................................................................3-4 Standard Weights .......................................................3-4 Draining the Fuel ........................................................3-4 Oil ...............................................................................3-4 Other Fluids ................................................................3-4 Configuration of the Aircraf ......................................3-5 Jacking the Aircraft ....................................................3-5 Leveling the Aircraft ..................................................3-5 Safety Considerations .................................................3-5 Determining the CG .......................................................3-5 EWCG Formulas............................................................3-6 Datum Forward of the Airplane—Nosewheel Landing Gear ..............................................................3-6 Datum Aft of the Main Wheels—Nosewheel Landing Gear ..............................................................3-6 Location of Datum .........................................................3-7 Datum Forward of the Main Wheels—Tailwheel Landing Gear ..............................................................3-7 Datum Aft of the Main Wheels—Tailwheel Landing Gear ..............................................................3-7 Center of Gravity (CG) and Mean Aerodynamic Chord (MAC) .................................................................3-8 Table of Contents x Center of Gravity (CG)...............................................3-8 Mean Aerodynamic Chord (MAC) ............................3-8 Chapter 4 Light Sport Aircraft Weight and Balance Control ..................................................................4-1 Introduction ....................................................................4-1 LSA Definition of Term ............................................4-2 Weight and Balance .......................................................4-3 WSC Aircraft..............................................................4-3 Powered Parachutes....................................................4-4 Weight and Balance Computations (Amateur-Built Aircraft)..........................................................................4-4 Chapter 5 Single-Engine Aircraft Weight and Balance Computations ......................................................5-1 Introduction ....................................................................5-1 Determining the Loaded Weight and CG ......................5-2 Manual Computational Method..................................5-2 Loading Graph Method ..............................................5-4 Moment Indexes .....................................................5-4 Loading Graph ........................................................5-4 Compute Weight and Balance Using the Loading Graph ........................................................5-4 Chapter 6 Multiengine Aircraft Weight and Balance Computations ......................................................6-1 Introduction ....................................................................6-1 Determine the Loaded CG .............................................6-2 Chart Method Using Weight, Arm, and Moments .....6-2 Determining the CG in Percentage of Mean Aerodynamic Chord (MAC) ......................................6-3 The Chart Method Using Weight and Moment Indexes .........................................................6-3 Chapter 7 Center of Gravity Change After a Repair or Alteration .............................................7-1 Introduction ....................................................................7-1 Equipment List ...............................................................7-2 Major Alteration and Repair ..........................................7-3 Weight and Balance Revision Record ...........................7-3 Weight Changes Caused by a Repair or Alteration .......7-5 Computations Using Weight, Arm, and Moment.......7-5 Computations Using Weight and Moment Indexes ...7-5 Determining the CG in Percentage of Mean Aerodynamic Chord (Percent MAC) .........................7-6 Empty Weight CG (EWCG) Range ...............................7-6 Adverse-Load CG Checks .............................................7-6 Forward Adverse-Load CG Check .............................7-7 Aft Adverse-Load CG Check .....................................7-7 Ballast ............................................................................7-7 Temporary Ballast ......................................................7-8 Temporary Ballast Formula........................................7-8 Permanent Ballast .......................................................7-8 Chapter 8 Weight and Balance Control—Helicopter .........8-1 Introduction ....................................................................8-1 Determining the Loaded CG of a Helicopter .................8-3 Effects of Offloading Passengers and Using Fue ..........8-3 Chapter 9 Weight and Balance Control—Commuter Category and Large Aircraft ...............................9-1 Introduction ....................................................................9-1 Establishing the Initial Weight of an Aircraft ...............9-2 Determining the Empty Weight and Empty Weight CG (EWCG)...................................................................9-2 Documenting Changes to an Aircraft’s Weight and Balance ..........................................................................9-3 Determining the Loaded CG of the Airplane in Percent MAC .................................................................9-3 Operational Empty Weight (OEW) ...............................9-5 Reestablishing the OEW ............................................9-5 Fleet Operating Empty Weights (FOEW) .................9-5 Onboard Aircraft Weighing System ...........................9-5 Determining the Correct Stabilizer Trim Setting .......9-6 Determining CG Changes Caused by Modifying the Cargo ..............................................................................9-6 Effects of Loading or Offloading Cargo ....................9-6 Effects of Shifting Cargo From One Hold to Another .......................................................................9-8 Determining Cargo Pallet Loads and Floor Loading Limits ...........................................................9-9 Determining the Maximum Amount of Payload That Can Be Carried .................................................9-10 Determining the Landing Weight .............................9-10 Determining Fuel Dump Time in Minutes ...............9-12 Weight and Balance of Commuter Category Airplanes ..................................................................9-12 Determining the Loaded Weight and CG .................9-12 Determining the Changes in CG When Passengers Are Shifted ...............................................................9-16 Determining Changes in Weight and CG When the Aircraft Is Operated in Its Cargo Configuratio ............................................................9-18 Determining the CG Shift When Cargo Is Moved From One Section to Another ..................................9-18 Determining the CG Shift When Cargo Is Added or Removed ..............................................................9-18 Determining Which Limits Are Exceeded ...............9-18 xi Chapter 10 Use of Computer for Weight and Balance Computations ....................................................10-1 Introduction ..................................................................10-1 Electronic Calculator ...................................................10-2 E6-B Flight Computer..................................................10-2 Dedicated Electronic Flight Computer ........................10-3 Typical Weight and Balance Problems ....................10-4 Determining CG in Inches From the Datum ............10-4 Nosewheel Airplane With Datum Ahead of the Main Wheels .........................................................10-4 Nosewheel Airplane With Datum Behind the Main Wheels .........................................................10-4 Tailwheel Airplane With Datum Ahead of the Main Wheels .........................................................10-4 Tailwheel Airplane With Datum Behind the Main Wheels .........................................................10-4 Determining CG, Given Weights, and Arms ...........10-5 Determining CG, Given Weights, and Moment Indexes......................................................................10-5 Determining CG in Percent Mean Aerodynamic Chord (MAC) ...........................................................10-5 Determining Lateral CG of a Helicopter ..................10-5 Determining ΔCG Caused by Shifting Weights.......10-6 Determining Weight Shifted to Cause Specified ΔCG ..........................................................................10-6 Determining Distance Weight Is Shifted to Move CG a Specific Distanc .............................................10-6 Determining Total Weight of an Aircraft With a Specified ΔCG When Cargo Is Move ....................10-7 Determining Amount of Ballast Needed to Move CG to a Desired Location .........................................10-7 Appendix A.......................................................... A-1 Appendix B.......................................................... B-1 Glossary ..............................................................G-1 Index ......................................................................I-1 xii 1-1 Introduction There are many factors in the safe and efficient operation of aircraft, including proper weight and balance control. The weight and balance system commonly employed among aircraft consists of three equally important elements: the weighing of the aircraft, the maintaining of the weight and balance records, and the proper loading of the aircraft. An inaccuracy in any one of these elements defeats the purpose of the system. The final loading calculations are meaningless if either the aircraft has been improperly weighed or the records contain an error. Improper loading decreases the effic ency and performance of an aircraft from the standpoint of altitude, maneuverability, rate of climb, and speed. It may even be the cause of failure to complete the flight or, for that matter, failure to start the flight. Because of abnormal stresses placed upon the structure of an improperly loaded aircraft, or because of changed fly ng characteristics of the aircraft, loss of life and destruction of valuable equipment may result. Weight and Balance Control Chapter 1 1-2 Aircraft can perform safely and achieve their designed efficiency only when they are operated and maintained in the way their designers intended. This safety and efficiency is determined to a large degree by holding the aircraft’s weight and balance parameters within the limits specified for its design. The remainder of this handbook describes how this is done. Responsibility for Weight and Balance Control The responsibility for proper weight and balance control begins with the engineers and designers and extends to the technicians who maintain the aircraft and the pilots who operate them. Modern aircraft are engineered utilizing state- of-the-art technology and materials to achieve maximum reliability and performance for the intended category. As much care and expertise must be exercised in operating and maintaining these efficient aircraft as was taken in their design and manufacturing: 1. The designers of an aircraft set the maximum weight based on the amount of lift the wings or rotors can provide under the operational conditions for which the aircraft is designed. The structural strength of the aircraft also limits the maximum weight the aircraft can safely carry. The designers carefully determine the ideal center of gravity (CG) and calculate the maximum allowable deviation from this specific location. 2. The manufacturer provides the aircraft operator with the empty weight of the aircraft and the location of its empty weight center of gravity (EWCG) at the time the certified aircraft leaves the factory. Amateur-built aircraft must have this information determined and available at the time of certification 3. The FAA-certificated mechanic or repairman who maintains the aircraft keeps the weight and balance records current, recording any changes that have been made because of repairs or alterations. 4. The pilot in command (PIC) has the responsibility prior to every flight to know the maximum allowable weight of the aircraft and its CG limits. This allows the pilot to determine during the preflight inspection that the aircraft is loaded so that the CG is within the allowable limits. Terminology Pilots and FAA-certificated mechanics or repairmen must ensure they understand the terms as they relate to the aircraft in question. For small aircraft terminology, use the information found in sources associated with Civil Air Regulation (CAR) 3 certification or General Aviation Manufacturers Association (GAMA) Specification No. 1 for part 23 aircraft or part 27 for rotorcraft. For terminology applied to large part 25 aircraft, information can be found in Advisory Circular (AC) 120-27, Aircraft Weight and Balance Control. The glossary contains the most current terms and defin tions. Current regulations are available from the Superintendent of Documents; U.S. Government Printing Office; Washington, DC 20402. They are also located on the FAA website at www.faa.gov. Earlier regulations may be available in libraries or in the Federal Register. Weight Control Weight is a major factor in airplane construction and operation, and it demands respect from all pilots and particular diligence by all maintenance personnel. Excessive weight reduces the efficiency of an aircraft and the available safety margin if an emergency condition should arise. When an aircraft is designed, it is made as light as the required structural strength allows, and the wings or rotors are designed to support the maximum allowable weight. When the weight of an aircraft is increased, the wings or rotors must produce additional lift and the structure must support not only the additional static loads, but also the dynamic loads imposed by flight maneuvers. For example, the wings of a 3,000-pound airplane must support 3,000 pounds in level flight, but when the airplane is turned smoothly and sharply using a bank angle of 60°, the dynamic load requires the wings to support twice this or 6,000 pounds. Severe uncoordinated maneuvers or flight into turbulence can impose dynamic loads on the structure great enough to cause failure. In accordance with Title 14 of the Code of Federal Regulations (14 CFR) part 23, the structure of a normal category airplane must be strong enough to sustain a load factor of 3.8 times its weight. Every pound of weight added to a normal category aircraft requires that the structure be strong enough to support 3.8 pounds. An aircraft operated in the utility category must sustain a load factor of 4.4 times its weight, and acrobatic category aircraft must be strong enough to withstand 6.0 times their weight. The lift produced by a wing is determined by its airfoil shape, angle of attack, speed through the air, and air density. When an aircraft takes off from an airport with a high density altitude, it must accelerate to a speed faster than would be required at sea level to produce enough lift to allow takeoff; therefore, a longer takeoff run is necessary. The distance needed may be longer than the available runway. When operating from a high density altitude airport, the Pilot’s Operating Handbook (POH) or Airplane Flight Manual (AFM) must be consulted to determine the maximum weight allowed for the aircraft under the conditions of altitude, temperature, wind, and runway conditions. 1-3 Effects of Weight Most modern aircraft are so designed that, when all seats are occupied, the baggage compartment is full, and all fuel tanks are full, the aircraft is grossly overloaded. This type of design requires the pilot to give great consideration to the requirements of each specific flight. If maximum range is required, occupants or baggage must be left behind, or if the maximum load must be carried, the range, dictated by the amount of fuel on board, must be reduced. Overloading an aircraft can create a variety of problems: • The aircraft needs a higher takeoff speed, which results in a longer takeoff run. • Both the rate and angle of climb are reduced. • The service ceiling is lowered. • The cruising speed is reduced. • The cruising range is shortened. • Maneuverability is decreased. • A longer landing roll is required because the landing speed is higher. • Excessive loads are imposed on the structure, especially the landing gear. The POH or AFM includes tables or charts that give the pilot an indication of the performance expected for any weight. An important part of careful preflight planning includes a check of these charts to determine if the aircraft is loaded so the proposed flight can be safely made Weight Changes The maximum allowable weight for an aircraft is determined by design considerations. However, the maximum operational weight may be less than the maximum allowable weight due to such considerations as high density altitude or high-drag field conditions caused by wet grass or water on the runway. The maximum operational weight may also be limited by the departure or arrival airport’s runway length. One important preflight consideration is the distribution of the load in the aircraft. Loading the aircraft so the gross weight is less than the maximum allowable is not enough. This weight must be distributed to keep the CG within the limits specified in the POH or AFM If the CG is too far forward, a heavy passenger can be moved to one of the rear seats or baggage may be shifted from a forward baggage compartment to a rear compartment. If the CG is too far aft, passenger weight or baggage can be shifted forward. The fuel load should be balanced laterally. The pilot should pay special attention to the POH or AFM regarding the operation of the fuel system in order to keep the aircraft balanced in flight Weight and balance of a helicopter is far more critical than for an airplane. Some helicopters may be properly loaded for takeoff, but near the end of a long flight when the fuel tanks are almost empty, the CG may have shifted enough for the helicopter to be out of balance laterally or longitudinally. Before making any long flight, the CG with the fuel available for landing must be checked to ensure it is within the allowable range. Airplanes with tandem seating normally have a limitation requiring solo flight to be made from the front seat in some airplanes or the rear seat in others. Some of the smaller helicopters also require solo flight be made from a specific seat, either the right, left, or center. These seating limitations are noted by a placard, usually on the instrument panel, and they should be strictly followed. As an aircraft ages, its weight usually increases due to debris and dirt collecting in hard-to-reach locations and moisture absorbed in the cabin insulation. This increase in weight is normally small, but it can be determined only by accurately weighing the aircraft. Changes of fixed equipment may have a major effect upon the weight of the aircraft. Many aircraft are overloaded by the installation of extra radios or instruments. Fortunately, the replacement of older, heavy electronic equipment with newer, lighter types results in a weight reduction. This weight change, however helpful, can cause the CG to shift, which must be computed and annotated in the weight and balance record. Repairs and alterations are the major sources of weight changes. It is the responsibility of the FAA-certificate mechanic or repairman making any repair or alteration to know the weight and location of a change, to compute the CG, record the new empty weight and EWCG in the aircraft weight and balance record, and update the equipment lists. If the newly calculated EWCG should happen to fall outside the EWCG range, it is necessary to perform an adverse- loading check. This requires a forward and rearward adverse-loading check and a maximum weight check. These weight and balance extreme conditions represent the maximum forward and rearward CG position for the aircraft. An adverse-loading check is a deliberate attempt to load an aircraft in a manner that creates the most critical balance condition and still remains within the design CG limits of the aircraft. If any of the checks fall outside the loaded CG range, 1-4 Figure 1-1. Longitudinal forces acting on an airplane in flight. Lift Variable Nose-up force dependent upon airspeed Fixed Nose-down force independent of airspeed CG Figure 1-2. If the CG is too far aft at the low stall airspeed, there might not be enough elevator nose-down authority to get the nose down for recovery. Lift Insufficient elevator Nose-down force CG too far aft CG Figure 1-3. If the CG is too far forward, there is not enough elevator nose-up force to flare the airplane for landing. Lift Insufficient elevator Nose-up force CG too far forward CG the aircraft must be reconfigured or placarded to prevent the pilot from loading the aircraft improperly. It is sometimes possible to install a fixed ballast in order for the aircraft to operate again within the normal CG range. The FAA-certificated mechanic or repairman conducting an annual or condition inspection must ensure the weight and balance data in the aircraft records is current and accurate. It is the responsibility of the PIC to use the most current weight and balance data when operating the aircraft. Stability and Balance Control Balance control refers to the location of the CG of an aircraft. This is of primary importance to aircraft stability, which is a factor in flight safety. The CG is the point at which the total weight of the aircraft is assumed to be concentrated, and the CG must be located within specific limits for safe flight. Both lateral and longitudinal balance are important, but the prime concern is longitudinal balance; that is, the location of the CG along the longitudinal or lengthwise axis. An airplane is designed to have stability that allows it to be trimmed to maintain straight-and-level flight with hands off the controls. Longitudinal stability is maintained by ensuring the CG is slightly ahead of the center of lift. This produces a fixed nose-down force independent of the airspeed. This is balanced by a variable nose-up force, which is produced by a downward aerodynamic force on the horizontal tail surfaces that varies directly with the airspeed. [Figure 1-1] If a rising air current should cause the nose to pitch up, the airplane slows and the downward force on the tail decreases. The weight concentrated at the CG pulls the nose back down. If the nose should drop in flight, the airspeed increases and the increased downward tail load brings the nose back up to level flight As long as the CG is maintained within the allowable limits for its weight, the airplane has adequate longitudinal stability and control. If the CG is too far aft, it is too near the center of lift; the airplane is unstable and difficult to recover from a stall. [Figure 1-2] If the unstable airplane should enter a spin, the spin could become flat making recovery difficult or impossible. If the CG is too far forward, the downward tail load needs to be increased to maintain level flight. This increased tail load has the same effect as carrying additional weight; the aircraft must fly at a higher angle of attack and drag increases. A more serious problem caused by the CG being too far forward is the lack of sufficient elevator authority. At low takeoff speeds, the elevator might not produce enough nose-up force to rotate; on landing there may not be enough elevator force to flare the airplane. [Figure 1-3] Both takeoff and landing runs are lengthened if the CG is too far forward. The basic aircraft design is such that lateral symmetry is assumed to exist. For each item of weight added to the left of the center line of the aircraft (also known as buttock line zero or BL-0), there is generally an equal weight at a corresponding location on the right. The lateral balance can be upset by uneven fuel loading or burnoff. The position of the lateral CG is not normally computed for an airplane, but the pilot must be aware of 1-5 Figure 1-4. Lateral imbalance causes wing heaviness, which may be corrected by deflecting the aileron. The additional lift causes additional drag, and the airplane flies inefficiently. Empty Full Additional lift and drag Additional weight Figure 1-5. Fuel in the tanks of a swept-wing airplane affects both lateral and longitudinal balance. As fuel is used from an outboard tank, the CG shifts forward. Inboard fuel: nose heavy Outboard fuel: tail heavy the adverse effects that result from a laterally unbalanced condition. [Figure 1-4] This is corrected by using the aileron trim tab until enough fuel has been used from the tank on the heavy side to balance the airplane. The deflected trim tab deflects the aileron to produce additional lift on the heavy side, but it also produces additional drag, and the airplane flies inefficientl Helicopters are more often affected by lateral imbalance than airplanes. If a helicopter is loaded with heavy occupants and fuel on the same side, it could be out of balance enough to make it unsafe to fly. It is also possible that if external loads are carried in such a position that requires large lateral displacement of the cyclic control to maintain level fl ght, the fore-and-aft cyclic control effectiveness is limited. Swept-wing airplanes are more critical due to fuel imbalance because as the fuel is used from the outboard tanks, the CG shifts forward. As fuel is used from the inboard tanks, the CG shifts aft. [Figure 1-5] For this reason, fuel-use scheduling in swept-wing airplanes operation is critical. Weight Control for Aircraft Other Than Fixed and Rotor-wing Some light aircraft utilize different methods of determining weight and balance from the traditional fixed and rotor-wing aircraft. These aircraft achieve flight control by methods different from the fixed-wing airplane or helicopter. Most notable of these are weight-shift control (WSC) aircraft (also known as trikes), powered parachutes, and balloons. These aircraft typically do not specify either an EWCG or a CG range. They require only a certified or approved maximum weight. To understand why this is so, a look at how flight control is achieved is helpful. Airplanes and WSC aircraft control flight under the influenc of the same four forces (lift, gravity, thrust, and drag), and around the same three axes (pitch, yaw, and roll). However, each aircraft accomplishes this control in a very different manner. This difference helps explain why the fixed-wing airplane requires an established weight and a known CG, whereas the WSC aircraft only requires the known weight. The fixed-wing airplane has movable controls that alter lift on various airfoil surfaces to vary pitch, roll, and yaw. In turn, these changes in lift affect the characteristics of the flight parameters. Weight normally decreases in flight due to fuel consumption, and the airplane CG changes with this weight reduction. An airplane utilizes its variable fligh controls to compensate and maintain controllability through the various flight modes and as the CG changes. An airplane has a CG range or envelope within which it must remain if the flight controls are to remain effective and the airplane safely operated. Weight-Shift Control Aircraft The WSC aircraft has a relatively set platform wing without a tail. The pilot achieves control by shifting weight. In the design of this aircraft, the weight of the airframe and its payload is attached to the wing at a single point in a pendulous arrangement. The pilot, through the flight controls, controls the arm of this pendulum and thereby controls the aircraft. When a change in flight parameter is desired, the pilot displaces the aircraft’s weight by the appropriate distance and direction. This change momentarily disrupts the equilibrium between the four forces acting on the aircraft. The wing, due to its inherent stability, then moves appropriately to reestablish the desired relationship between these forces; the wing flexes and alter its shape. As the shape is changed, lift is varied at different points on the wing to achieve the desired flight parameters 1-6 The flight controls primarily affect the pitch-and-roll axes. Since there is no vertical tail plane, there is minimal or no yaw control. Unlike in an airplane, the CG experienced by the WSC aircraft wing remains constant. Since the weight of the airframe acts through a single point (the wing attach point), the range over which the weight may act is fixed at the pendulum arm or length. Even though weight decreases as fuel is consumed, weight remains focused at the wing attach point. Because the range is fixed, there is no need to establish a calculated range. The pilot should consult the POH or Aircraft Operating Instructions (AOI) for maximum takeoff weight and minimum and maximum seat weight limits prior to each flight Powered Parachutes The powered parachute is also a pendulum-style aircraft. Its airframe CG is fixed at the pendulum attach point. It is more limited in controllability than the WSC aircraft because it lacks an aerodynamic pitch control. Pitch (and lift) control is primarily a function of the power control. Increased power results in increased lift; cruise power amounts to level flight; decreased power causes a descent. Due to this characteristic, the aircraft is basically a one-airspeed aircraft. Once again, because the CG is fixed at the attach point to the wing, there is no CG range. As with WSC, the pilot should consult the POH or AOI for maximum takeoff weight and minimum and maximum seat weight limits prior to each flight Roll control on a powered parachute is achieved by changing the shape of the wing. The change is achieved by varying the length of steering lines attached to the outboard trailing edges of the wing. The trailing edge of the parachute is pulled down slightly on one side or the other to create increased drag along that side. This change in drag creates roll and yaw, permitting the aircraft to be steered. Balloons The balloon is controlled by the pilot only in the vertical dimension; this is in contrast to all other aircraft. He or she achieves this control through the use of lift and weight. Wind provides all other movement. The CG of the gondola remains constant beneath the balloon envelope. As in WSC and powered-parachute aircraft, there is no CG limitation. Underestimating the Importance of Weight and Balance Many pilots, from sport pilot to commercial pilot, tend to underestimate the importance of proper weight and balance of their aircraft. Load sheets are taken for granted and hasty calculations are made of the aircraft’s CG. Unfortunately, each year there are a number of accidents related to weight and balance issues. Many of these occurrences could have been avoided had more attention been given to weight and balance. Every student pilot is taught how to work a weight and balance problem and that it is important to make sure every flight is loaded “within the envelope” (no more than maximum gross weight) for both takeoff and landing. But does he or she really understand just why this is so and the disastrous effect of being out of the envelope? Two examples of documented cases are provided below in an effort to indicate the serious nature of maintaining the proper weight and balance. In case studies when weight and balance was listed as the major factor of the accident, many were fatal. For instance, a small aircraft was loaded with hunters, gear, and dogs (none of the dogs were secured inside the aircraft). During takeoff, all the dogs went to the aft of the airplane. This shifted the CG well aft of its allowable limit. The airplane stalled and crashed. The airplane was destroyed with casualties. Another accident occurred when a group of skydivers were sitting on the floor toward the aft portion of the airplane (they were unsecured). During takeoff, the CG was again well beyond its aft limit. The airplane stalled and crashed. The airplane was destroyed with casualties. There is a safety factor built into the formula for maximum gross weight. Any airplane can fly when it takes off at a weight greater than maximum gross weight if the runway is long enough and the density altitude is low enough. However, landing is a different matter. All airplanes are built to withstand an occasional hard landing, but what would happen if the hard landing were combined with a substantially overweight airplane? Something would probably break at that time or the structure would be weakened enough to break sometime in the future when everything might seem normal to a pilot unaware of the previous situation. Even more disastrous than an overweight, hard landing is reaching or exceeding the structural integrity of the metal and/or composite design values when maneuvering or when turbulence is encountered. Hidden damage could result, causing an unexpected catastrophic failure at some future time. If an airplane is certificated with a maximum gross weight of 6,000 pounds (its weight on the ground) and is rolled into a 60° bank, the forces exerted make it feel as if it weighed 12,000 pounds. At its maximum certificated gross weight, there is no problem because the aircraft is operated within its certificated maneuvering loads. But loaded to 8,000 pounds with a 60° bank or an abrupt pullup, it suddenly weighs 16,000 pounds and might not be able to perform! Even if it could, there would probably be internal stress damage that would show up on future flights 2-1 Introduction Weight and balance in aircraft is based on the law of the lever. This chapter discusses the application of the law of the lever and its applications relative to locating the balance point of a beam or lever on which various weights are located or shifted. The chapter also discusses the documentation pertaining to weight and balance that is furnished by the Federal Aviation Administration (FAA) and aircraft manufacturers. Weight and Balance Theory Chapter 2 2-2 Figure 2-1. Balance lever. B = 200 A = 100 Arm B (+25") Arm A (−50") Fulcrum Datum − Moment + Moment − F o r c e s + F o r c e s Figure 2-2. Balance point locations. Item Weight (lb) Arm (in) Moment (lb-in) −5,000 +5,000 0 Weight A Weight B 100 200 300 −50 +25 Figure 2-3. Balance lever datum located off the lever. C = 200 A = 100 B = 100 +50 +90 +110 +150 CG Datum Weight and Balance Theory Two elements are vital in the weight and balance considerations of an aircraft. • The total weight of the aircraft must be no greater than the maximum weight allowed by the FAA for the make and model of the aircraft. • The center of gravity (CG), or the point at which all of the weight of the aircraft is considered to be concentrated, must be maintained within the allowable range for the operational weight of the aircraft. Arm The arm is usually measured and expressed in inches and refers to the horizontal distance between the CG of an item or object and the datum, a point from where all measurements are taken. Arms to the left of the datum are negative (–) and those to the right of the datum are positive (+). The datum is an imaginary vertical plane from which all horizontal distances are measured for balance purposes. The position of the reference datum varies by aircraft design and manufacturer. When the datum is located off of the lever and to the left, all of the arms are positive and computational errors are minimized. Note: When the datum is established ahead of the aircraft, for example at the aircraft nose, all of the arms are positive and computational errors are minimized. Moment A moment is a force that tries to cause rotation and is the product of the arm, in inches, and the weight, in pounds. Moments are generally expressed in pound-inches (lb-in) and may be either positive or negative. The Law of the Lever Weight and balance problems are based on the physical law of the lever. This law states that a lever is balanced when the weight on one side of the fulcrum (a pivot point for the lever) multiplied by its arm is equal to the weight on the opposite side multiplied by its arm. In other words, the lever is balanced when the sum of the moments about the fulcrum is zero. This is the condition in which the positive moments (those that try to rotate the lever clockwise) are equal to the negative moments (those that try to rotate it counterclockwise). In an aircraft, the balance point is referred to as the CG. One of the easiest ways to understand weight and balance is to consider a lever with weights placed at various locations. The balance point or CG of the lever can be changed by either moving the weights closer or farther from the fulcrum or by increasing or decreasing the weights. The balance point or CG of a lever may be determined by using these four steps: 1. Measure the arm of each weight in inches from the datum. 2. Multiply each arm by its weight in pounds to determine the moment in pound-inches of each weight. 3. Determine the total of all weights and of all the moments. (Disregard the weight of the lever). 4. Divide the total moment by the total weight to determine the balance point. Consider these facts about the lever in Figure 2-1. The 100-pound weight A is located 50 inches to the left of the fulcrum (the datum, in this instance), and it has a moment of 100 × –50 = –5,000 lb-in. The 200-pound weight B is located 25 inches to the right of the fulcrum, and its moment is 200 × +25 = +5,000 lb-in. In Figure 2-2, the sum of the moments is –5,000 + 5,000 = 0, and the lever is balanced. The forces that try to rotate it clockwise have the same magnitude as those that try to rotate it counterclockwise. If either weight is moved or changed, the balance point or CG changes and the lever becomes unbalanced. In Figure 2-3, the datum is located off the lever to the left of weight A. Using the information provided in Figure 2-3, 2-3 Figure 2-4. Finding balance point with datum located off the lever. Item Weight (lb) Arm (in) Moment CG 5,000 9,000 30,000 44,000 110 Weight A Weight B Weight C 100 100 200 400 50 90 150 Figure 2-5. Locating balance point. C = 200 A = 100 B = 100 −20 −60 110 +40 CG Original datum New datum Figure 2-6. Proving balance point with three weights is correct. Item Weight (lb) Arm (in) Moment (lb-in) −6,000 −2,000 +8,000 0 Weight A Weight B Weight C 100 100 200 −60 −20 +40 Figure 2-7. Locating balance point with datum at C. C = 200 A = 100 B = 100 −60 −100 Datum Figure 2-8. Determining new balance point. Item Weight (lb) Arm (in) Moment CG −10,000 −6,000 0 −16,000 −40 Weight A Weight B Weight C 100 100 200 400 −100 −60 Figure 2-9. Locating balance point with datum left of original. C = 200 A = 100 B = 100 −40 Datum −20 −60 +40 determine the balance point by making a chart like the one in Figure 2-4. As noted in Figure 2-4, A weighs 100 pounds and is 50 inches from the datum; B weighs 100 pounds and is 90 inches from the datum; C weighs 200 pounds and is 150 inches from the datum. The total of the weights is 400 pounds, and the total moment is 44,000 lb-in. Determine the balance point by dividing the total moment by the total weight. A balance point is equal to the CG and can be mathematically written as: CG = total moment total weight To prove this is the correct balance, move the datum to a location 110 inches to the right of the original datum and determine the arm of each weight from this new datum. [Figure 2-5] Then, make a new chart similar to the one in Figure 2-6. If the balance point is correct, the sum of the moments is zero. The new arm of weight A is 60 inches (the difference between 110 and 50), and since this weight is to the left of the datum, its arm is negative or –60 inches. The new arm of weight B is 20 inches (110 – 90), and it is also to the left of the datum, so it is –20; the new arm of weight C is 40 inches (150 – 110). It is to the right of the datum and is therefore positive. The lever is balanced when the sum of the moments is zero. The location of the datum used for determining the arms of the weights is not important; it may be in various locations, but all of the measurements must be made from the same datum location. The procedure for finding the balance point is the same anywhere the datum is located. In Figure 2-7, the datum is located at C. Weight A has an arm of –100 inches (negative because it is to the left) of the datum and weight B has an arm of –60 inches from the datum. The table in Figure 2-8 is used to determine the new balance point. To verify that this is the correct balance point, move the datum 40 inches to the left of the original datum and determine the arm of each weight from this new datum as in Figure 2-9. 2-4 Figure 2-10. Proving the new balance point is correct. Item Weight (lb) Arm (in) Moment (lb-in) −6,000 −2,000 +8,000 0 Weight A Weight B Weight C 100 100 200 −60 −20 40 Figure 2-11. Locating balance point with three weights. C = 200 B = 200 A = 100 80 72 100 CG Datum Before Weight Shift Figure 2-12. Proving the new balance point is correct. Item Weight (lb) Arm (in) Moment (lb-in) −5,000 +10,000 +5,000 Weight A Weight B Weight C 100 200 −50 +50 Figure 2-13. Weight distribution to balance lever. B = 200 A = 100 C = 100 −50 50 −25 CG Datum After Weight Shift Figure 2-14. Weight shift provides correct CG. Item Weight (lb) Arm (in) Moment (lb-in) −5,000 −5,000 +10,000 0 Weight A Weight B Weight C 100 200 200 −50 −25 +50 The new arm for weight A would be –100 + 40 = –60; for weight B, –60 + 40 = –20; and point C, is +40. The lever is balanced and the balance point is correct when the sum of the moments is zero. [Figure 2-10] Shifting the Balance Point or CG One common weight and balance problem involves moving or shifting weight from one point to another in order to move the balance point or CG to a desired location. This can be demonstrated by using a lever with three weights to work out the problem. Solution by Chart As the lever is loaded in Figure 2-11, it balances at a point 72 inches from the CG of weight A. To shift weight B so the lever balances about its center, 50 inches from the CG of weight A, first determine the arm of weight B that produces a moment that causes the total moment of all three weights around this desired balance point to be zero. The combined moment of weights A and C around this new balance point is 5,000 lb-in, so the moment of weight B must be –5,000 lb-in for the lever to balance. [Figure 2-12] Determine the arm of weight B by dividing its moment, –5,000 lb-in, by its weight of 200 pounds. The arm is –25 inches. To balance the lever at its center, weight B must be placed so its CG is 25 inches to the left of the center of the lever. [Figure 2-13] Figure 2-14 indicates that the shift in weight depicted in Figure 2-13 allows the lever to balance as the sum of the moments is zero. Basic Weight and Balance Equation The following formulas can be used to determine the distance weight must be shifted to obtain a desired change in the CG location. The equation can also be rearranged to fi d the amount of weight required to be shifted to move the CG to a desired location, to find the distance the CG is moved when a specified amount of weight is shifted, or to find the total weight that would allow shifting a specified amount of weight to move the CG a given distance. Weight to be shifted = Δ CG Total weight Distance weight is shifted Total weight = Weight shifted × Distance weight is shifted Δ CG Weight shifted = Total weight shifted × Δ CG Distance weight is shifted Δ CG = Weight shifted × Distance weight is shifted Total weight Distance weight is shifted = Total weight × Δ CG Weight shifted 2-5 Figure 2-14. Center of gravity expressed as percent mean aerodynamic chord. MAC MAC 30% MAC 15% MAC Datum Quarter chord Half semi-span Solution by Formula The problem in Figure 2-11 can be solved by using variations of this basic equation. First, rearrange the formula to determine the distance weight B must be shifted: Distance weight B is shifted = Total weight × Δ CG Weight shifted = 500 × –22 200 = –55 inches The CG of the lever in Figure 2-11 was 72 inches from the datum. This CG can be shifted to the center of the lever as in Figure 2-13 by moving weight B. If the 200-pound weight B is moved 55 inches to the left, the CG shifts from +72 inches to +50 inches, a distance of 22 inches. When the distance the weight is to be shifted is known, the amount of weight to be shifted to move the CG to any location can be determined by another arrangement of the basic equation. Use the following arrangement of the formula to determine the amount of weight that has to be shifted from station 8 to station +25, to move the CG from station +72 to station +50. Weight shifted = Total weight × Δ CG Distance weight is shifted = 500 × 22 55 = 200 inches If the 200-pound weight B is shifted from station +80 to station +25, the CG moves from station +72 to station +50. A third arrangement of this basic equation is used to determine the amount the CG is shifted when a given amount of weight is moved for a specified distance (as it was done in Figure 2-11). The following formula is used to determine the amount the CG is shifted when 200-pound weight B is moved from +80 to +25. Δ CG = Weight shifted × Distance it is shifted Total weight = 200 × 55 500 = 22 inches Moving weight B from +80 to +25 moves the CG 22 inches from its original location at +72 to its new location at +50 as seen in Figure 2-13. To complete the calculations, return to the original formula and enter the appropriate numbers. Weight to be shifted = Δ CG Total weight Distance weight is shifted 200 = 22 500 55 .4 = .4 The equation is balanced. Mean Aerodynamic Chord The CG point affects the stability of the aircraft. To ensure the aircraft is safe to fly, the CG must fall within specified limits established by the manufacturer. On some aircraft, the CG is expressed as a percentage of the length of the mean aerodynamic chord (MAC) or “percent MAC.” [Figure 2-14] In order to make such a calculation, the position of the leading edge of the MAC must be known ahead of time. CG limits are specified forward and aft and/or lateral (left and right) limits within which the aircraft’s CG must be located during flight. The area between the limits is called the CG range of the aircraft. 2-6 The position of the fore and aft CG limits is measured as a percentage of MAC from the MAC leading edge. Usually for a single or two-seat aircraft, the most forward position would be forward of 30–35 percent MAC. Thus, the allowable CG range in a light aircraft should not exceed 20 percent MAC. Note: For a rectangular width of constant airfoil section dimensions, MAC is just the chord. For a symmetrically- tapered wing, it is the average of the root chord and the tip chord. FAA-Furnished Weight and Balance Information The information discussed to this point can be readily applied to any aircraft weight and balance problem. To apply the techniques, certain elements of information are required. This information is obtained from both FAA documents and manufacturer provided data. Before an aircraft CG can be computed, certain information must be known. This information, furnished by the FAA for every certificated aircraft in the Type Certificate Data Sheets (TCDS) or Aircraft Specifications, can be accessed at www.faa.gov. When the design of an aircraft is approved by the FAA, an Approved Type Certificate and TCDS are issued. The TCDS includes all of the pertinent specification for the aircraft; at each annual or 100-hour inspection, it is the responsibility of the inspecting mechanic or repairman to ensure that the aircraft adheres to them. A note about the TCDS: aircraft certifi ated before January 1, 1958, were issued Aircraft Specifica ions under the Civil Air Regulations (CARs), but when the Civil Aeronautical Administration (CAA) was replaced by the FAA, Aircraft Specification were replaced by the TCDS. The weight and balance information on a TCDS includes CG range, empty weight CG range (EWCG), maximum weights, number of seats, maximum baggage, fuel capacity, oil capacity, and datum location. Data pertinent to an individual model is located in its respective section of the TCDS. Use the TCDS excerpts in Figure 2-15 to verify the following for a Cirrus Design Corporation SR 20 normal category: CG Range S/N 1005 through 1147: Forward Limits: 138.7 inches at 2,110 lb with a straight line taper to 141.0 in at 2,694 lb and 143.0 in at 2,900 lb Aft Limits: 144.6 in at 2,110 lb, with straight line taper to 147.4 in at 2,570 lb, and to 147.9 in at 2,745 lb, and 148.2 in at 2,900 lb S/N 1148 through 1877, 1879 through 1885, and S/N 1005 through 1147 if Cirrus Service Bulletin SB 20-01-00 is complied with: Forward Limits: 138.7 in at 2,110 lb with a straight line taper to 141.0 in at 2,694 lb and 144.1 in at 3,000 lb Aft Limits: 144.6 in at 2,110 lb, with straight line taper to 147.4 in at 2,570 lb, and to 148.1 in at 2,900 lb, and 148.0 in at 3,000 lb S/N 1878, 1886 and Subsequent: Forward Limits: 137.8 in at 2,100 lb with a straight line taper to 139.1 in at 2,700 lb, and to 140.7 in at 3,050 lb Aft Limits: 148.1 in at 2,100 lb, with straight line to 148.1 in at 3,050 lb Empty Weight CG Range (EWCG) When all of the seats and baggage compartments are located close together, it is not possible (as long as the EWCG is located within the EWCG range) to legally load the aircraft so that its operational CG falls outside this allowable range. If the seats and baggage areas extend over a wide range, the EWCG range is listed as “None.” Maximum Weights The maximum allowable takeoff and landing weights and the maximum allowable ramp weight are given. This basic information may be altered by a note. Notes are found in data pertinent to all models. An example would be Note 6 at the end of Figure 2-15. Number of Seats The number of seats and their arms are given in such terms as: 4 (2 at 143.5 aft of datum, 2 at 180 aft of datum). Maximum Baggage Maximum baggage for this model is 130 pounds at 208 inches. Fuel Capacity This important information is given in such terms as: 60.5 gal at 153.75 in. Usable: 56 gal (See Note 1). Notes can be found in data pertinent to all models. Oil Capacity (Wet Sump) The quantity of the full oil supply and its arm are given: 8 quarts at 76.2 in. 2-7 Figure 2-15. Sample excerpt from TCDS A00009CH. P a g e N o . 1 2 3 4 5 6 7 R e v . N o . 1 1 1 2 1 2 1 4 1 4 1 6 1 6 DEPARTMENT OF TRANSPORTATION FEDERAL AVIATION ADMINISTRATION A00009CH Revision 16 Cirrus Design Corporation SR20 SR22 April 22, 2009 TYPE CERTIFICATE DATA SHEET NO. A00009CH This data sheet, which is part of Type Certificate No. A00009CH, prescribes conditions and limitations under which the product for the which type certificate was issued meets the airworthiness requirements of the Federal Aviation Regulations. Type Certificate Holder: Cirrus Design Corporation 4515 Taylor Circle Duluth, MN 55811 I - Model SR20, (Normal Category), Approved October 23, 1998 Engine Teledyne Continental IO-360-ES, Type Certificate Data Sheet (TCDS) E1CE Fuel 100/100LL minimum grade aviation gasoline Engine Limits Maximum Take-off 2700 RPM (200 hp) Maximum Continuous Power 2700 RPM (200 hp) Propeller and Propeller limits 1. Hartzell Propeller Inc. P/N BHC-J2YF-1BF/F7694 TCDS P37EA Maximum Diameter: 76 inches Minimum Diameter: 73 inches Number of Blades: 2 Low Pitch: 14.6°+/-0.5° High Pitch: 35.0°+/-1.0° Not to be operated above 24 inches of manifold pressure between 1900 and 2200 RPM. Spinner: Hartzell P/N A-2295(P) NOTE: Spinner may be painted or polished. 2. Hartzell Propeller Inc. P/N PHC-J3YF-1MF/F7392-1 TCDS P36EA Maximum Diameter: 74 inches Minimum Diameter: 72 inches Number of Blades: 3 Low Pitch: 14.1°+/-0.5° High Pitch: 35.0°+/-1.0° No operating limitations to 2800 RPM Spinner: Hartzell P/N A-2295-1P 3. Hartzell Propeller Inc. P/N PHC-J3YF-1RF/F7392-1 TCDS P36EA Maximum Diameter: 74 inches Minimum Diameter: 72 inches Number of Blades: 3 Low Pitch: 13.9°+/-0.5° High Pitch: 35.0°+/-1.0° No operating limitations to 2800 RPM Spinner: Hartzell P/N A-2295-1(P) NOTE: Spinner may be painted or polished. 2-8 Figure 2-15. Sample excerpt from TCDS A00009CH (continued). A 0 0 0 0 9 C H - 2- Airspeed Limits S/N 1005 thru 1147: Vne Never Exceed Speed 200 KIAS Vno Maximum Structural Cruising Speed 165 KIAS Vo (2900 lbs) Operating Maneuvering Speed 135 KIAS Vo (2600 lbs) Operating Maneuvering Speed 126 KIAS Vo (2200 lbs) Operating Maneuvering Speed 116 KIAS Vfe Maximum Flap Extension Speed 100 KIAS Vpd Maximum Parachute Deployment Speed 135 KIAS S/N 1148 thru 1877, 1879 thru 1885, and S/N 1005 thru 1147 if Cirrus Service Bulletin SB 20-01- 00 is complied with: Vne Never Exceed Speed 200 KIAS Vno Maximum Structural Cruising Speed 165 KIAS Vo (3000 lbs) Operating Maneuvering Speed 131 KIAS Vo (2600 lbs) Operating Maneuvering Speed 122 KIAS Vo (2300 lbs) Operating Maneuvering Speed 114 KIAS Vfe Maximum Flap Extension Speed 100 KIAS Vpd Maximum Parachute Deployment Speed 135 KIAS S/N 1878, 1886 and subsequent: Vne Never Exceed Speed 200 KIAS Vno Maximum Structural Cruising Speed 163 KIAS Vo (3050 lbs) Operating Maneuvering Speed 130 KIAS Vfe Maximum Flap Extension Speed 104 KIAS Vpd Maximum Parachute Deployment Speed 133 KIAS C.G. Range S/N 1005 thru 1147: Forward Limits: 138.7 inches at 2110 lbs with a straight line taper to 141.0 inches at 2694 lbs, and 143.0 inches at 2900 lbs. Aft Limits: 144.6 inches at 2110 lbs, with straight line taper to 147.4 inches at 2570 lbs, and to 147.9 inches at 2745 lbs, and 148.2 inches at 2900 lbs. S/N 1148 thru 1877, 1879 thru 1885, and S/N 1005 thru 1147 if Cirrus Service Bulletin SB 20-01- 00 is complied with: Forward Limits: 138.7 inches at 2110 lbs with a straight line taper to 141.0 inches at 2694 lbs, and 144.1 inches at 3000 lbs. Aft Limits: 144.6 inches at 2110 lbs, with straight line taper to 147.4 inches at 2570 lbs, and to 148.1 inches at 2900 lbs, and 148.0 inches at 3000 lbs. S/N 1878, 1886 and subsequent: Forward Limits: 137.8 inches at 2100 lbs with a straight line taper to 139.1 inches at 2700 lbs, and to 140.7 inches at 3050 lbs Aft Limits: 148.1 inches at 2100 lbs, with straight line to 148.1 inches at 3050 lbs. Empty Weight C.G. Range None Maximum Weight S/N 1005 thru 1147: Takeoff and Landing: 2900 lbs. S/N 1148 thru 1877, 1879 thru 1885, and S/N 1005 thru 1147 if Cirrus Service Bulletin SB 20-01- 00 is complied with: Takeoff: 3000 lbs. Landing: 2900 lbs. Zero Fuel: 2900 lbs. S/N 1878, 1886 and subsequent: Takeoff and Landing: 3050 lbs. 2-9 Figure 2-15. Sample excerpt from TCDS A00009CH (continued). Figure 2-15. Sample excerpt from TCDS A00009CH (continued). - 3- A 0 0 0 0 9 C H Minimum Crew One (1) Pilot Number of Seats 4 (2 at 143.5 inches aft of datum, 2 at 180 inches aft of datum) Maximum Baggage 130 Lbs. at 208 inches Fuel Capacity Total: S/N 1005 thru 1877, 1879 thru 1885 60.5 gal at 153.75 inches Usable: 56 gal (See Note 1) S/N 1878, 1886 and subsequent: 58.5 gal at 154.9 inches Usable: 56 gal (See Note 1) Oil Capacity 8 quarts at 76.2 inches Maximum Operating Altitude With a portable oxygen system, the aircraft is limited to 17,500 ft MSL. Oxygen must be provided as required by the operating rules. Only portable oxygen systems listed in the FAA Approved Airplane Flight Manual, document number 11934-002, or later FAA approved revisions, are allowed. Control Surface Movements Wing Flaps: Up 0°± 0.5° Down 50% 16°± 0.5° Down 100% 32°± 0.5° Aileron: Up 12.5° ± 1.0° Down 12.5° ±1.0° Elevator: Up 25.0° +0°/-1.0° Down 15° ± 1.0° Elevator Trim: Up 17.0° Minimum Down 10.5° ± 1.0° Rudder: Right 20.0° ± 1.0° Left 20.0° ± 1.0° Additional Limitations: Airframe life limit: 12,000 flight hours Design Data: The airplane shall be manufactured in accordance with the latest FAA approved revision of “Master Drawing List”, Document No. 13750, or other FAA approved data. NOTE: Document No. 12609 is the predecessor document to Document No. 13750. Serial Nos. Eligible 1005 and on II - Model SR22, Normal Category, Approved November 30, 2000 Engine Teledyne Continental IO-550-N, Type Certificate Data Sheet E3SO Engine Limits Maximum Take-off 2700 RPM (310 hp) Maximum Continuous Power 2700 RPM (310 hp) Propeller and Propeller limits 1. Hartzell Propeller Inc. P/N PHC-J3YF-1RF/F7694 or F7694B TCDS P36EA Hartzell Maximum Diameter: 78 inches Minimum Diameter: 76 inches Number of Blades: 3 Low Pitch: 14.1°+/-0.5° High Pitch: 35.0°+/-1.0° No operating limitations to 2700 RPM Spinner: Hartzell P/N A-2295-1(P) NOTE: Spinner may be painted or polished. 2. McCauley Propeller Systems P/N D3A34C443/78CYA-0 TCDS P47GL McCauley Maximum Diameter: 78 inches Minimum Diameter: 76 inches Number of Blades: 3 - 5- A 0 0 0 0 9 C H Minimum Crew One (1) Pilot Number of Seats 4 (2 at 143.5 inches aft of datum, 2 at 180 inches aft of datum) Maximum Baggage 130 Lbs. at 208 inches Fuel Capacity Total: S/N 0002 thru 2333, 2335 thru 2419, and 2421 thru 2437 84 gallon at 154.9 inches Usable: 81 gallon (See Note 1) S/N 2334, 2420, 2438 and subsequent 94.5 gallon at 154.9 inches Usable: 92.0 gallon (See Note 1) or 58.5 gal at 154.9 inches Usable: 56 gal (See Note 1) Oil Capacity 8 quarts at 77.1 inches Maximum Operating Altitude With a portable oxygen system, the aircraft is limited to 17,500 ft MSL. Oxygen must be provided as required by the operating rules. Only portable oxygen systems listed in the FAA Approved Airplane Flight Manual, document number 13772-001, or later FAA approved revisions, are allowed. Control Surface Movements Wing Flaps: Up 0°±0.5° Down 50% 16° ± 0.5° Down 100% 32° ±0.5° Aileron: Up 12.5° ± 1.0° Down 12.5° ± 1.0° Aileron Trim: Up 6 ± 1.0° Down 6 ± 1.0° Elevator: Up 25.0° +0°/-1.0° Down 15° ± 1.0° Elevator Trim: Up 17.0° Minimum Down 10.5° ± 1.0° Rudder: Right 20.0° ± 1.0° Left 20.0° ± 1.0° Additional Limitations: Airframe life limit: 12,000 flight hours Design Data: The airplane shall be manufactured in accordance with the latest FAA approved revision of “Master Drawing List”, Document No. 13750, or other FAA approved data. Serial Nos. Eligible 0001 and on. Data Pertinent to All Models Reference Datum 100 inches in front of the forward face of firewall bulkhead Leveling Means Door sill and leveling points as defined in AFM Certification Basis Model SR20: 14 CFR Part 23 of the Federal Aviation Regulations effective February 1, 1965, as amended by 23-1 through 23-47, except as follows: 14 CFR 23.573, 23.575, 23.611, 23.657, 23.673 through Amendment 23-48; 14 CFR 23.783, 23.785, 23.867, 23.1303, 23.1307, 23.1309, 23.1311, 23.1321, 23.1323, 23.1329, 23,1361, 23.1383, 23.1401, 23.1431, 23.1435 through Amendment 23-49; 14 CFR 23.3, 23.25, 23.143, 23.145, 23.155, 23.1325, 23.1521, 23.1543, 23.1555, 23.1559, 23.1567, 23.1583, 23.1585, 23.1589 through Amendment 23-50; 14 CFR 23.777, 23.779, 23.901, 23.907, 23.955, 23.959, 23.963, 23.965, 23.973, 23.975, 23.1041, 23.1091, 23.1093, 23.1107, 23.1121, 23.1141, 23.1143, 23.1181, 23.1191, 23.1337 through Amendment 23-51; 14 CFR 23.1305 through Amendment 23-52 Noise: 14 CFR Part 36 dated December 1, 1969 as amended by 36-1 through 36-21. 2-10 Figure 2-16. Weight and balance report. Item C.G. Arm (in) Moment (lb-in/1,000) Standard empty weight Optional equipment Special installation Paint Unusable fuel Basic empty weight 1,876.0 1.2 6.2 — 30.0 1,913.4 67,798.6 16.7 257.3 — 1,380.0 69,452.6 36.1 13.9 41.5 — 46.0 Aircraft Serial No. 18259080 FAA Registration No. N42565 Date: 4-22-09 Weight (lb) × Figure 2-15. Sample excerpt from TCDS A00009CH (continued). Note 2. All placards specified in the latest FAA approved revisions of the "PILOT'S OPERATING HANDBOOK AND FAA APPROVED AIRPLANE FLIGHT MANUAL FOR THE CIRRUS SR20", document numbers 11934-001, 11934-002, 11934-003 or 11934-004 and the latest FAA approved revisions of the "PILOT'S OPERATING HANDBOOK AND FAA APPROVED AIRPLANE FLIGHT MANUAL FOR THE CIRRUS SR22" document numbers 13772-001 or 13772-002 must be displayed in the airplane in the appropriate locations. Note 3. FAA approved Airworthiness Limitations are included in Section 4 of the Airplane Maintenance Manual (AMM) Document No. 12137-001 for model SR20, and 13773-001 for model SR22. Note 4. Exterior colors are limited to those specified in the latest FAA accepted revision of the Airplane Maintenance Manual (AMM) Document No. 12137-001 for model SR20, and 13773-001 for model SR22. Note 5. Major structural repairs must be accomplished in accordance with FAA approved Cirrus Design repair methods or other methods approved by the FAA. Note 6. For Model SR22 S/N 0002 thru 2333, 2335 thru 2419, and 2421 thru 2437 a maximum landing weight exists along the line between 141.4 inches at 3210 lbs and 142.7 inches at 3400 lbs. -- END -- Data Pertinent to all Models The location of the datum is specified and is described, for example, as: 100 inches in front of the forward face of the firewall bulkhead. Manufacturer-Furnished Information When an aircraft is initially certificated, its empty weight and EWCG are determined and recorded in the weight and balance record, such as the one in Figure 2-16. Notice in this figur that the moment is expressed as “Moment (lb-in/1,000).” This is a moment index, which means that the moment, a very large number, has been divided by 1,000 to make it more manageable. Chapter 4, Light Sport Aircraft—Weight and Balance Control, discusses moment indices in more detail. The aircraft is furnished with an equipment list, specifies all the required equipment and all equipment approved for installation in the aircraft. The weight and arm of all installed equipment are included on the list and checked prior to the aircraft leaving the factory. When an aircraft mechanic or repairman adds or removes any item on the equipment list, he or she must change the weight and balance record to indicate the new empty weight and EWCG, and the equipment list is revised to show which equipment is actually installed. Figure 2-17 is from a comprehensive equipment list that includes all of the items of equipment approved for this particular model of aircraft. The pilot’s operating handbook (POH) for each individual aircraft includes an aircraft-specific equipment list of the items from this master list. When any item is added to or removed from the aircraft, its weight and arm are determined in the equipment list and used to update the weight and balance record. The POH and airplane flight manual (AFM) also contain CG moment envelopes and loading graphs. Examples of the use of these helpful graphs are given in Chapter 5, Single-Engine Aircraft Weight and Balance Computations. In addition to the weight and balance report and equipment list, the manufacturer also provides a CG range chart. The CG range can be found in text form in the TCDS. The CG range chart is furnished in the POH. 2-11 Figure 2-17. Excerpt from a typical comprehensive equipment list. 2-12 Figure 2-17. Excerpt from a typical comprehensive equipment list (continued). 3-1 Introduction Chapter 2, Weight and Balance Theory, explained the theory of weight and balance and gave examples of the way the center of gravity (CG) could be found for a lever loaded with several weights. In this chapter, the practical aspects of weighing an airplane and locating its CG are discussed. Formulas are introduced that allow the CG location to be measured in inches from various datum locations and in percentage of the mean aerodynamic chord (MAC). Weighing the Aircraft and Determining the Empty Weight Center of Gravity Chapter 3 3-2 Figure 3-1. Platform scales. Requirements Regulations do not require periodic weighing of privately owned and operated aircraft. Such aircraft are usually weighed when originally certificated or after major alterations that can affect the weight and balance. The primary purpose of aircraft weight and balance control is safety. Manufacturers conduct extensive flight tests to establish loading limits for their aircraft because limit information is critical for safe flight. A secondary purpose is to aid efficiency during flight. Overloading of the aircraft is not the only concern; the distribution of the weight is important also. The aircraft has CG limits, and any loading that places the CG outside the established limits seriously impairs controllability of the aircraft. Weight and balance is of such vital importance that each Federal Aviation Administration (FAA) certificated mechanic or repairman maintaining an aircraft must be fully aware of his or her responsibility to provide the pilot with current and accurate information for the actual weight of the aircraft and the location of the CG. The pilot in command (PIC) is responsible for knowing the weight of the load, CG, maximum allowable weight, and CG limits of the aircraft. The weight and balance report must include an equipment list showing weights and moment arms of all required and optional items of equipment included in the certificate empty weight. Weight and balance records used in accounting for and correcting the CG location are reliable for only limited periods of time. For this reason, periodic aircraft weighing is desirable. An aircraft should be reweighed and a new weight and balance record should be started after the aircraft has undergone extensive repair or major alteration, when the pilot reports unsatisfactory flight characteristics (e.g., nose or tail heaviness), and when recorded weight and balance data are suspected to be in error. Repairs and alterations are major sources of weight change. The airframe and powerplant (A&P) FAA-certificated mechanic or repairman who is responsible for making any repair or alteration must: 1. Establish by computation that the authorized weight and CG limits as shown in the type certificate data sheet (TCDS) and aircraft specifications are not exceeded, and 2. Record the new empty weight center of gravity (EWCG) data in the current approved aircraft flight manual or issued operating limitations. When an aircraft has undergone extensive repair or major alteration, it should be reweighed and a new weight and balance record started. The A&P FAA-certificated mechanic or repairman responsible for the work must provide the pilot with current and accurate aircraft weight information and location of the EWCG. Equipment for Weighing Weighing aircraft with accurately calibrated scales is the only sure method of obtaining an accurate empty weight and CG location. The two basic types of scales used to weigh aircraft are platform and load cell. Platform scales [Figure 3-1] or ramp wheel scales [Figure 3-2] (usually a form or modified version of the platform scale) are low profi e, easy to handle, safe, and reliable. Tow or push the aircraft wheels or skids onto the scale pad at ground level. With one scale per wheel, each device should be capable of measuring up to at least 60,000 pounds since the weight on each wheel rarely exceeds this figure Load cell scales [Figure 3-3] are also a reliable means to weigh aircraft and are typically cheaper than the platform type.Using load cell scales allows for the aircraft to be set up and weighed in its level flight attitude. With this method, the aircraft is placed on jacks with electronic load cells placed between the jack and the jack pad on the aircraft. The aircraft is raised on the jacks until the wheels or skids are off the floor and the aircraft is in a level flight attitude. The weight measured by each load cell is indicated on the control panel. Jacking an aircraft off the ground from all load points can be an inconvenience, as well as a safety risk, which some operators would rather avoid by opting for more expensive—but simpler to use—platform equipment. In addition, weighing with platform scales typically takes only one-third of the time needed to weigh with load cells. 3-3 Figure 3-2. Ramp scales. Figure 3-3. Load cell scales. All scales for aviation use, manual or electronic, must be protected when stored or shipped, and they must be checked periodically for accuracy. The maximum recognized period between calibration checks is 12 months; however, this period may be reduced by an airworthiness authority dependent on the conditions of use. Scales in daily use may require a shorter interval and/or testing to determine the continued accuracy of the unit. Scales should be returned to the manufacturer for proper calibration and testing. Preparation for Weighing In general, weight procedures may vary with the aircraft and types of weight equipment employed. The weighing procedure contained in the manufacturer’s maintenance manual should be followed for each particular aircraft. The major considerations in preparing an aircraft for weighing are described in the following paragraphs. Scale Preparation Mechanical and electronic scales shall be inspected prior to use and set to zero. This is done by adding and removing a weight, then rechecking for zero. This process should be repeated until a steady zero setting is obtained. The scales should be located in the same environment in which they 3-4 Figure 3-4. Standard fuels and weights with temperatures of 32 °F and 59 °F. Weighing Point AVGAS (Aviation Gasoline) JET A & A-1 Water Oil 6.14 6.75 8.35 7.50 6.01 6.68 8.33 7.43 Pounds per U.S. Gallon 32 °F 59 °F are to be used and allowed to come up to temperature at least 2 hours prior to use. Scales should not be used in temperature extremes below 40 °F or above 100 °F unless the scale is specifically designed for use in those temperatures. Electronic scales are very sensitive and, if subjected to freezing temperatures, the liquid displays may be damaged beyond use. Weigh Clean Aircraft Inside Hangar The aircraft should be weighed inside a hangar where wind cannot blow over the surface and cause fluctuating or false scale readings. The aircraft should be clean inside and out, with special attention paid to the bilge area to ensure that no water or debris is trapped there. The outside of the aircraft should be as free as possible of all mud and dirt. Equipment List All of the required equipment must be properly installed, and there should be no equipment installed that is not included in the equipment list. If such equipment is installed, the weight and balance record must be corrected to indicate it. Ballast All required permanent ballasts must be properly secured in place. All temporary ballasts must be removed. Standard Weights Standard weights are established weights for numerous items involved in weight and balance computations. These weights should not be used if actual weights are available. Some of the standard weights are listed in Figure 3-4. Note the difference in weight as temperatures change. Although this change is a very small amount per gallon, it could end up in a significant total weight gain when dealing with large quantities of fluids, such as those found in commercial aircraft. Draining the Fuel Drain fuel from the tanks in the manner specified by the aircraft manufacturer. If there are no specific instructions, drain the fuel until the fuel quantity gauges read empty when the aircraft is in level-flight attitude. Any fuel remaining in the system is considered residual or unusable fuel and is part of the aircraft empty weight. The amount of residual fuel and its arm are normally found in Note 1 in the section of the Type Certificate Data Sheets (TCDS), “Data pertaining to all Models.” For additional fuel capacity information, see Chapter 2, Weight and Balance Theory. If it is not feasible to drain the fuel, the tanks can be topped off to be sure of the quantity they contain and the aircraft weighed with full fuel. After weighing is complete, the weight of the fuel and its moment are subtracted from those of the aircraft as weighed. To correct the empty weight for the residual fuel, add its weight and moment. When computing the weight of the fuel (e.g., a tank full of jet fuel), measure its specific gravity (sg) with a hydrometer and multiply it by 8.345 (the nominal weight of 1 gallon of pure water whose sg is 1.0). If the ambient temperature is high and the jet fuel in the tank is hot enough for its specific gravity to reach 0.81 rather than its nominal sg of 0.82, the fuel actually weighs 6.76 pounds per gallon rather than its normal weight of 6.84 pounds per gallon. Oil The empty weight for aircraft certificated under the Civilian Air Regulations (CAR) part 3 does not include the engine lubricating oil. The oil must either be drained before the aircraft is weighed, or its weight must be subtracted from the scale readings to determine the empty weight. To weigh an aircraft that does not include the engine lubricating oil as part of the empty weight, place it in level flight attitude, then open the drain valves and allow the oil to drain out. Any remaining is undrainable oil and is part of the empty weight. Aircraft certificated under Title 14 of the Code of Federal Regulations (14 CFR) parts 23 and 25 include full oil as part of the empty weight. If it is impractical to drain the oil, the reservoir can be filled to the specified level and the weight of the oil computed at 7.5 pounds per gallon. Then, its weight and moment are subtracted from the weight and moment of the aircraft as weighed. The amount and arm of the undrainable oil are found in Note 1 of the TCDS, and this must be added to the empty weight. Other Fluids The hydraulic fluid reservoir and all other reservoirs containing fluids required for normal operation of the aircraft should be full. Fluids not considered to be part of the empty weight of the aircraft are potable (drinkable) water, lavatory precharge water, and water for injection into the engines. 3-5 Figure 3-5. Locating the CG of an airplane relative to the datum. Weighing point Net weight (lb) TARE weight (lb) Scale reading (lb) Arm (in) Moment (lb-in) CG 46.0 46.0 −32.0 32.8 Right side Left side Nose Total 16 16 8 830 836 340 2,006 38,180 38,456 −10,880 65,756 846 852 348 Configuration of the Aircraf Consult the aircraft service manual regarding position of the landing gear shock struts and the control surfaces for weighing. When weighing a helicopter, the main rotor must be in its correct position. Jacking the Aircraft Aircraft are often weighed by rolling them onto ramps in which load cells are embedded. This eliminates the problems associated with jacking the aircraft off the ground. However, many aircraft are weighed by jacking the aircraft up and then lowering them onto scales or load cells. Extra care must be used when raising an aircraft on jacks for weighing. If the aircraft has spring steel landing gear and it is jacked at the wheel, the landing gear will slide inward as the weight is taken off of the tire, and care must be taken to prevent the jack from tipping over. For some aircraft, stress panels or plates must be installed before the aircraft is raised with wing jacks to distribute the weight over the jack pad. Be sure to follow the recommendations of the aircraft manufacturer in detail anytime an aircraft is jacked. When using two wing jacks, take special care to raise them simultaneously, keeping the aircraft so it does not slip off the jacks. As the jacks are raised, keep the safety collars screwed down against the jack cylinder to prevent the aircraft from tilting if one of the jacks should lose hydraulic pressure. Leveling the Aircraft When an aircraft is weighed, it must be in its level fligh attitude so that all of the components are at their correct distance from the datum. This attitude is determined by information in the TCDS. Some aircraft require a plumb line to be dropped from a specified location so that the point of the weight (the bob) hangs directly above an identifiable point. Others specify that a spirit level be placed across two leveling lugs, often special screws on the outside of the fuselage. Other aircraft call for a spirit level to be placed on the upper door sill. Lateral level is not specified for all light aircraft, but provisions are normally made on helicopters for determining both longitudinal and lateral level. This may be done by built-in leveling indicators, or by a plumb bob that shows the conditions of both longitudinal and lateral level. The actual adjustments to level the aircraft using load cells are made with the jacks. When weighing from the wheels, leveling is normally done by adjusting the air pressure in the nosewheel shock strut. Safety Considerations Special precautions must be taken when raising an aircraft on jacks. 1. Stress plates must be installed under the jack pads if the manufacturer specifies them 2. If anyone is required to be in the aircraft while it is being jacked, there must be no movement. 3. The jacks must be straight under the jack pads before beginning to raise the aircraft. 4. All jacks must be raised simultaneously and the safety devices are against the jack cylinder to prevent the aircraft tipping if any jack should lose pressure. Not all jacks have screw down collars, some use drop pins or friction locks. Determining the CG When the aircraft is in its level flight attitude, drop a plumb line from the datum and make a mark on the hangar floor below the tip of the bob. Draw a chalk line through this point parallel to the longitudinal axis of the aircraft. Then, draw lateral lines between the actual weighing points for the main wheels, and make a mark along the longitudinal line at the weighing point for the nosewheel or the tailwheel. These lines and marks on the floor allow accurate measurements between the datum and the weighting points to determine their arms. Determine the CG by adding the weight and moment of each weighing point to determine the total weight and total moment. Then, divide the total moment by the total weight to determine the CG relative to the datum. As an example of locating the CG with respect to the datum, which in this case is the firewall, consider the tricycle landing gear airplane as detailed in the Figure 3-5 table and illustrated in Figure 3-6. When the airplane is on the scales with the parking brakes 3-6 Figure 3-9. Determining the CG with datum forward of an airplane with nosewheel landing gear. CG F × L W = D − = 128 − 114.8 = ( ) 340 × 78 2,006 ( ) Figure 3-8. The datum is 100 inches forward of the wing root leading edge. D = 128.0 100.0 13.2 50 114.8 Datum L = 78.0 Figure 3-10. The datum is aft of the main wheels at the wing trailing edge. 153 D = 75 88.2 Datum L = 78.0 Figure 3-7. Determining the CG. CG Total moment Total weight = 65,756 2,006 = 32.8 inches behind the datum = Figure 3-6. The datum is located at the firewall. 46.0 32 32.8 Datum off, place chocks around the wheels to keep the airplane from rolling. Subtract the weight of the chocks, called tare weight, from the scale reading to determine the net weight at each weighing point. Multiply each net weight by its arm to determine its moment, and then determine the total weight and total moment. The CG is determined by dividing the total moment by the total weight. [Figure 3-7] The airplane illustrated in Figures 3-5 and 3-6 has a net weight of 2,006 pounds, and its CG is 32.8 inches behind the datum. EWCG Formulas A chart such as the one in Figure 3-5 helps the pilot visualize the weights, arms, and moments when solving an EWCG problem, but it is quicker to determine the EWCG by using formulas and an electronic calculator. The use of a calculator for solving these problems is described in Chapter 8, Use of Computers in Weight and Balance Computations. There are four possible conditions and their formulas that relate the location of CG to the datum. Notice that the formula for each condition first determines the moment of the nosewheel or tailwheel and then divides it by the total weight of the airplane. The arm thus determined is then added to or subtracted from the distance between the main wheels and the datum, distance D. Datum Forward of the Airplane—Nosewheel Landing Gear The datum of the airplane in Figure 3-8 is 100 inches forward of the leading edge of the wing root or 128 inches forward of the main-wheel weighing points. This is distance (D). The weight of the nosewheel (F) is 340 pounds, and the distance between main wheels and nosewheel (L) is 78 inches. The total weight of the airplane (W) is 2,006 pounds. Determine the CG by using the formula in Figure 3-9. The CG is 114.8 inches aft of the datum. This is 13.2 inches forward of the main-wheel weighing points, which proves the location of the datum has no effect on the location of the CG if all measurements are made from the same location. Datum Aft of the Main Wheels—Nosewheel Landing Gear The datum of some aircraft may be located aft of the main wheels. The airplane in this example is the same one just discussed, but the datum is at the intersection of the trailing edge of the wing with the fuselage. The distance (D) between the datum of the airplane in Figure 3-10 and the main-wheel 3-7 Figure 3-13. Determining the CG with datum forward of the main wheels in an airplane with tailwheel landing gear. CG R × L W = D + = 7.5 + 19.7 = ( ) 67 × 222 1,218 ( ) Figure 3-11. Determining the CG with datum aft of the main wheels of an airplane with nosewheel landing gear. CG F × L W = − D + = − 75 + −88.2 = ( ) 340 × 78 2,006 ( ) Figure 3-12. The datum of this tailwheel airplane is the wing root leading edge. D = 7.5 Datum 229.5 L = 222.0 19.7 Datum L = 222.0 D = −80 142 67.8 Figure 3-14. The datum is aft of the main wheels, at the intersection of the wing trailing edge and the fuselage. CG R × L W = −D + = −80 + = −67.8 ( ) 67 × 222 1,218 ( ) Figure 3-16. Determining the CG with datum aft of the main wheels in an airplane with tailwheel landing gear. weighing points is 75 inches, the weight of the nosewheel (F) is 340 pounds, and the distance between main wheels and nosewheel (L) is 78 inches. The total net weight of the airplane (W) is 2,006 pounds. The location of the CG may be determined by using the formula in Figure 3-11. The CG location is a negative value, which means it is 88.2 inches forward of the datum. This places it 13.2 inches forward of the main wheels, exactly the same location as when it was measured from other datum locations. Location of Datum The location of the datum is not important, but all measurements must be made from the same location. Datum Forward of the Main Wheels—Tailwheel Landing Gear Locating the CG of a tailwheel airplane is done in the same way as locating it for a nosewheel airplane except the formula is . The distance (D) between the datum of the airplane in Figure 3-12 and the main-gear weighing points is 7.5 inches, the weight of the tailwheel (R) is 67 pounds, and the distance (L) between the main-wheel and the tailwheel weighing points is 222 inches. The total weight of the airplane (W) is 1,218 pounds. Determine the CG by using the formula in Figure 3-13. R × L W The CG is 19.7 inches behind the datum. Datum Aft of the Main Wheels—Tailwheel Landing Gear The datum of the airplane in Figure 3-14 is located at the intersection of the wing root trailing edge and the fuselage. This places the arm of the main gear (D) at –80 inches. The net weight of the tailwheel (R) is 67 pounds, the distance between the main wheels and the tailwheel (L) is 222 inches, and the total net weight (W) of the airplane is 1,218 pounds. Since the datum is aft of the main wheels, use the formula found in Figure 3-15. The CG is 67.8 inches forward of the datum or 12.2 inches aft of the main-gear weighing points. The CG is in exactly the same location relative to the main wheels, regardless of 3-8 Figure 3-16. The MAC is the chord drawn through the geographic center of the plan area of the wing. CG MAC Trailing edge mean aerodynamic chord Leading edge mean aerodynamic chord TEMAC LEMAC Datum where the datum is located. Center of Gravity (CG) and Mean Aerodynamic Chord (MAC) Center of Gravity (CG) In addition to overloading or drastically reducing the aircraft’s weight, the distribution of weight is also a concern. When aircraft equipment is changed, the person making the equipment change must make an entry on the equipment list indicating items added, removed, or relocated; the date of the change; and the person’s name and certification number in the aircraft’s maintenance records. For the purpose of weight and balance computations, the CG of an airplane is an imaginary point about which the nose- heavy (–) moments and tail-heavy (+) moments are exactly equal in magnitude. If suspended from the CG point, the aircraft would have no tendency to rotate nose up or nose down. The CG of the loaded aircraft can range fore and aft within certain limits that are determined during the fl ght test for type certific tion. These limits are the most forward- and rearward-loaded CG positions at which the aircraft meets the performance and flight characteristics required by the FAA. Any loading that places the CG outside the limits for a particular aircraft seriously impairs the pilot’s ability to control the aircraft. For example, it is more difficult to take off and gain altitude in a nose-heavy aircraft, and the aircraft tends to drop its nose when the pilot reduces throttle. It also requires a higher speed to land safely. A tail-heavy aircraft is more susceptible to stalling at low speed, which is a concern during the landing approach. On small airplanes and on all helicopters, the CG location is identified as being a specific number of inches from the datum. The CG range is identified in the same way. On larger airplanes, from private business jets to large jumbo jets, the CG and its range are typically identified in relation to the width of the wing. Mean Aerodynamic Chord (MAC) The width of the wing, or straight-line distance from the leading edge to the trailing edge, on an airplane is known as the chord. If the leading edge and the trailing edge of a wing are parallel, the chord is equal at all points along the entire length of the wing. The average length of the chord, or MAC, of a tapered wing is more complicated to define. The MAC, as seen in Figure 3-16, is the chord of an imaginary airfoil that has the same aerodynamic characteristics as the actual airfoil. You can also think of it as the chord drawn through the geographic center of the plan area of the wing. Usually listed in the aircraft’s TCDS when it is required for weight and balance computations, the MAC is established by the manufacturer, defining its leading edge (LEMAC) and trailing edge (TEMAC) in terms of inches from the datum. [Figure 3-16] The CG location and various limits are then expressed in percentage of the chord or percent MAC. In order to relate the percent MAC to the datum, all weight and balance information includes two items: the MAC length in inches and the location of the LEMAC in inches from the datum. For the purpose of simplicity, most light- aircraft manufacturers express the CG range in inches from the datum; transport-category aircraft CGs are expressed in percent MAC. The relative positions of the CG and the aerodynamic center of lift of the wing have critical effects on the flight characteristics of the aircraft. Consequently, relating the CG location to the chord of the wing is convenient from a design and operations standpoint. Normally, an aircraft has acceptable fl ght characteristics if the CG is located somewhere near the 25 percent average chord point. This means the CG is located one-fourth of the distance back from the LEMAC to the TEMAC. Such a location places the CG forward of the aerodynamic center for most airfoils. The weight and balance data of the airplane in Figure 3-17 states that the MAC is from stations 144 to 206 and the CG is located at station 161. MAC = TEMAC – LEMAC MAC = 206" – 144" MAC = 62" 3-9 Figure 3-18. Formula for determining the CG expressed in percent MAC. CG inches % MAC = Distance aft of LEMAC × 100 MAC = = 27.4 17 × 100 62 Figure 3-17. Large aircraft weight and balance calculation diagram. CG = 27.4% CG = 161 MAC TEMAC LEMAC Datum CG = 27.4% CG CG 1 = 1611 61 MAC Station 144" Station 206" Figure 3-19. Formula for determining the CG in inches from the datum. CG inches from datum = LEMAC + MAC × CG % MAC 100 = 144 + = 160.9 62 × 27.4 100 CG = 161" In order to find the percent MAC, firs determine the distance of the CG from LEMAC. CG – LEMAC = Distance from LEMAC 161" – 144" = 17" The location of the CG expressed in percent MAC is determined using the formula found in Figure 3-18. The CG of the airplane is located at 27.4 percent MAC. It is sometimes necessary to determine the location of the CG in inches from the datum when its location in percent MAC is known. The CG of the airplane is located at 27.4 percent MAC. MAC = 206 – 144 = 62 LEMAC = station 144 Determine the location of the CG in inches from the datum by using the formula found in Figure 3-19. The CG of this airplane is located at station 160.9 inches aft of the datum. It is important for longitudinal stability that the CG be located ahead of the center of lift of a wing. Since the center of lift is expressed as percent MAC, the location of the CG is expressed in the same terms. 3-10 4-1 Introduction This chapter discusses the weight and balance procedures for light sport aircraft (LSA) that differ from conventional aircraft, specifically weight-shift control (WSC) aircraft (also called trikes), powered parachutes, and amateur-built LSA. [Figure 4-1] Light Sport Aircraft Weight and Balance Control Chapter 4 4-2 Figure 4-1. Examples of light sport aircraft (from top left, clockwise): weight-shift control, powered parachute, glider, airplanes, hot air balloon, and amateur-built LSA. LSA Definition of Term LSA is a category of simple, very basic, small, light-weight, low-performance aircraft, other than a helicopter or powered- lift, and a classification of aircraft specific to the United States. The Federal Aviation Administration (FAA) defines LSA as an aircraft with a maximum gross takeoff weight of not more than 1,320 pounds (600 kg) for aircraft not intended for operation over water, or 1,430 pounds (650 kg) for aircraft intended for operation over water; a maximum airspeed in level flight of 120 knots (220 kilometers per hour (km/h); 140 m