(d) If an acceptable method of analysis is developed by an applicant to statistically establish a satisfactory confidence level for the resulting parametric relationships, then 12 tests, in each aerodynamic configuration for which certification is desired, should be sufficient. More tests should be conducted if the distribution of the data does not give sufficient confidence in the parametric correlation. Past experience has shown that a total of 40 landings would establish a satisfactory confidence level without further analysis. Autolands may be included in the analysis but should not comprise more than half of the data points. If it is apparent that configuration is not a significant variable, all data may be included in a single parametric analysis.
(e) If an applicant proposes any other method as being equivalent to a parametric analysis, that method should be based on a developed mathematical model that employs performance-related variables such as thrust, attitude, angle-of-attack, and load factor to adequately reproduce the flight test trajectory and airspeed variation from the 50 foot point to touchdown. Such a mathematical model should be validated by not less than 12 tests in each aerodynamic configuration for which certification is desired, and be justified by a comparison of tested and calculated landing airborne distances.
(f) For the same aerodynamic configuration as previously certificated -- if new tests are necessary to substantiate performance to a weight higher than that permitted by the extrapolation limits of § 25.21(d), two landings per configuration should be conducted for each 5 percent increase in landing weight (but no more than a total of six landings should be needed).
These may be merged with previous certification tests for parametric analysis, whether the previous certification was conducted by this method or not. If a new aerodynamic configuration is proposed, the 12 tests per configuration described in paragraph (d), above, should be conducted.
(g) In calculating the AFM landing distances, the speed loss from 50 feet to touchdown, as a percentage of VREF, may be determined using the conditions described in paragraph 19b(3).
(4) Whichever method is chosen to establish airborne distances, satisfactory flight characteristics should be demonstrated in the flare maneuver when a final approach speed of VREF-5 knots is maintained down to 50 feet.
(a) Below 50 feet, the application of longitudinal control to initiate flare should occur at the same altitude as for a normal “on-speed” landing; no nose depression should be made and power should not be increased to facilitate the flare.
(b) All power levers should be in their minimum flight idle position prior to touchdown.
(c) The normal flare technique should be used such that the touchdown speed should be at least 5 knots less than the touchdown speed used to establish the landing distance and the rate of descent at touchdown should not be greater than 6 feet per second.
(d) This demonstration is performed at both maximum landing weight and near minimum landing weight.
(e) These V REF-5 knots landing demonstrations should not require the use of high control forces or full control deflections.
c. Procedures for Determination of the Transition and Stopping Distances .
(1) The transition distance extends from the initial touchdown point to the point where all approved deceleration devices are operative. The stopping distance extends from the end of transition to the point where the airplane is stopped. The two phases may be combined if the applicant prefers this method of analysis.
(2) If sufficient data are not available , there should be a minimum of six landings in the primary landing configuration. Experience has shown that if sufficient data are available for the airplane model to account for variation of braking performance with weight, lift, drag, ground speed, torque limit, etc., at least two test runs are necessary for each configuration when correlation for multiple configurations is being shown.
(3) A series of at least six measured landing tests covering the landing weight range should be conducted on the same set of wheels, tires, and brakes in order to substantiate that excessive wear of wheel brakes and tires is not produced in accordance with the provisions of § 25.125(c)(2). The landing tests should be conducted with the normal operating brake pressures for which the applicant desires approval. The main gear tire pressure should be set to not less than the maximum pressure desired for certification corresponding to the specific test weight.
Longitudinal control and brake application procedures should be such that they can be consistently applied in a manner that permits the airplane to be de-rotated at a controlled rate to preclude an excessive nose gear touchdown rate and so that the requirements of § 25.125(b)(4) and (5) are met. Nose gear touchdown rates in the certification landing tests should not be greater than eight feet per second. Certification practice has not allowed manually applied brakes before all main gear wheels are firmly on the ground.
(4) Describe the airplane operating procedures appropriate for determination of landing distance in the performance section of the AFM.
(5) Propeller pitch position used in determining the normal all-engines-operating landing stopping distance should be established using the criteria of § 25.125(g) for those airplanes that may derive some deceleration benefit from operating engines. Section 25.125(g) states that if the landing distance determined using a “device” that depends on the operation of any engine would be “noticeably increased” when a landing is made with that engine inoperative, the landing distance must be determined with that engine inoperative, unless a “compensating means” will result in one-engine-inoperative landing distances not greater than those with all engines operating. Acceptable interpretations of the terms “device,” “noticeably increased,” and “compensating means” are described below.
(a) If the propeller produces drag at any speed during the stopping phase of the normal all-engines-operating landing distance, the maximum drag from this “device” for which performance credit may be taken is that which results from a propeller pitch position that renders zero thrust at zero airspeed. If the normal operational ground idle setting produces negative thrust at zero airspeed, the all-engines-operating stopping distances should be determined using a special flight test power lever stop to limit the propeller blade angle.
(b) Distances should be measured for landings made with the propeller feathered on one engine, and ground idle selected after touchdown on the operating engines. The airplane configuration for this test, including the ground idle power lever position, should be the same as that used for the all-engines-operating landing distance determination. The nose wheel should be free to caster, as in VMCG tests, to simulate wet runway surface conditions. Differential braking may be used to maintain directional control. This testing should be conducted at a minimum of three weights that cover the expected range of operational landing weights. If the resulting distances do not exceed the all-engines-operating landing distances by more than two percent (2%), they are not “noticeably increased” and no further testing is required to take performance credit for all-engines-operating disking drag in AFM landing distances.
(c) If the distances determined in paragraph (b), above, are more than two percent greater than the all-engines-operating landing distances, there should be a “compensating means” in order to take performance credit for the all-engines-operating disking drag. Reverse propeller thrust on the operating engines is considered a “compensating means” if the resulting landing distances, with one propeller feathered, are demonstrated to be not longer than those determined for all-engines-operating with disking drag. The airplane configuration for this test should be the same as that used for the all-engines-operating landing distance determination. The nose wheel should be free to caster, as in VMCG tests, to simulate wet runway surface conditions.
Differential braking may be used to maintain directional control. Reverse thrust should not be selected until one second after nose wheel touchdown. This testing should be conducted at a minimum of three weights that cover the expected range of operational landing weights.
d. Instrumentation and Data . Instrumentation should include a means to record the airplane’s glide path relative to the ground, and the ground roll against time, in a manner that permits determining the horizontal and vertical distance time-histories. The appropriate data to permit analysis of these time-histories should also be recorded.
e. Landing on Unpaved Runways . Guidance material for evaluation of landing on unpaved runways is contained in Chapter 8 of this AC.
f. Automatic Braking Systems. Guidance material relative to evaluation of auto-brake systems is provided in paragraph 55c(6) of this AC.
g. Airplane Flight Manual Landing Distances .
(l) As a minimum, the AFM must include data for standard temperature and zero runway gradient showing the variation of landing distance with weight (up to maximum takeoff weight), altitude, and wind as required by §§ 25.1587(b) and 25.125(a). If the airplane is
78 intended for operation under part 121, the distances presented should include the operational field length factors for both dry and wet runways required by § 121.195.
(2) In accordance with § 25.101(i), AFM landing distances must be determined.
with all the airplane wheel brake assemblies at the fully worn limit of their allowable wear range.
The brakes may be in any wear state during the flight tests used to determine the landing distances, as long as a suitable combination of airplane and dynamometer tests is used to determine the landing distances corresponding to fully worn brakes. Alternatively, the relationship between brake wear and stopping performance established during accelerate-stop testing may be used if it encompasses the brake wear conditions and energies achieved during the airplane flight tests used to establish the landing distances.
(3) In deriving the scheduled distances, the time delays shown below should be assumed.
Figure 19-1. Landing Time Delays
(a) This segment represents the flight test measured average time from touchdown to pilot activation of the first deceleration device. For AFM data expansion, use the longer of 1 second or the test time.
(b) This segment represents the flight test measured average test time from pilot activation of the first deceleration device to pilot activation of the second deceleration device. For AFM data expansion, use the longer of 1 second or the test time.
(c) Step is repeated until pilot activation of all deceleration devices has been completed and the airplane is in the full braking configuration.
(4) For approved automatic deceleration devices (e.g., autobrakes or auto-spoilers, etc.) for which performance credit is sought for AFM data expansion, established times
79 determined during certification testing may be used without the application of the 1-second minimum time delay required in the appropriate segment above.
(5) It has been considered acceptable to expand the airborne portion of the landing distance in terms of a fixed airborne time, independent of airplane weight or approach speed.
(6) Assumptions to be made in assessing the effect of wind on landing distance are discussed in paragraph 3 of this AC.
h. Parametric Analysis Data Reduction . The following is an acceptable method of converting the test data to a mathematical model for the parametric analysis method of air distance described in paragraph 19b(3).
Test Data for Each Test Point:
R/S 50 = Rate of sink at 50 ft. above landing surface, Ft/Sec R/S TD = Rate of sink at touchdown, Ft/Sec V 50 = True airspeed at 50 ft. above landing surface, Ft/Sec V TD = True airspeed at touchdown, Ft/Sec t = Air time 50 ft. to touchdown, Sec
The multiple linear regression analysis as outlined below is used to solve for the constant of the two independent variable equations:
50/t = a + b(R/S 50) + (c)(R/STD)
To maintain the same units for all variables, the dependent variable is chosen as 50/t.
The test values of all the test points, 1 through n, are processed as follows, where n equals the number of test points and R1 through R13 are the regression coefficients:
R1 = 1n R/S50 R2 = 1n (R/S50)2 R3 = 1n R/STD R4 = 1n (R/STD)2 R5 = 1n (R/S50)(R/STD) R6 = 1n (50/t) R7 = 1n (R/S50)(50/t) R8 = 1n (R/STD)(50/t) R9 = (n)(R2)-(R1) 2
80 R10 = (n)(R8)-(R3)(R6) R11 = (n)(R5)-(R1)(R3) R12 = (n)(R7)-(R1)(R6) R13 = (n)(R4)-(R3) 2 c = ((R9)(R10)-(R11 )(R12))/((R9)(R13)-(R11)2) b = ((R12)-(c)(R11))/R9 a = ((R6)-(b)(R1)-(c)(R3))/n In the same manner, determine the values of the constants, a, b, and c, in an equation for speed reduction between 50 ft. and touchdown by replacing 50/t with (V50/VTD) for each test run.
After determining the values of the constants, the two equations are used to calculate the time from 50 ft. to touchdown and V50/VTD for the desired conditions of -3.5 degrees flight path and R/STD = 8 ft/Sec. The R/S50 is calculated from the approach path and V50.
After VTD is determined, the air distance may be determined for the average flare speed and t.
Example:
Test Data:
Run R/S 50 R/S TD V50 V TD t
1 13.4 6.1 219 214 5.6 2 10.9 1.8 223 218 8.5 3 7.9 5.8 209 201 7.4 4 8.3 2.3 213 206 9.6 5 9.8 4.1 218 212 7.5
Results:
50/t = 1.0432 + .3647(R/S 50) + .4917(R/STD)
V 50/VTD = 1.05508 - .003198(R/S50) + .001684(R/STD)
For conditions of V 50 = 220, flight path = -3.5 degrees, R/STD = 8.0, the results are:
R/S 50 = 13.43 V 50/VTD = 1.0256
t = 5.063 sec. Air Distance = 1100 ft.
Section 3. Controllability and Maneuverability
20. General - § 25.143.
a. Explanation . The purpose of § 25.143 is to verify that any operational maneuvers conducted within the operational envelope can be accomplished smoothly with average piloting skill and without encountering a stall warning or other characteristics that might interfere with normal maneuvering, or without exceeding any airplane structural limits. Control forces should not be so high that the pilot cannot safely maneuver the airplane. Also, the forces should not be so light that it would take exceptional skill to maneuver the airplane without over-stressing it or losing control. The airplane response to any control input should be predictable to the pilot.
(1) The maximum forces given in the table in § 25.143(d) for pitch and roll control for short term application are applicable to maneuvers in which the control force is only needed for a short period. Where the maneuver is such that the pilot will need to use one hand to operate other controls (such as during the landing flare or a go-around, or during changes of configuration or power resulting in a change of control force that needs to be trimmed out) the single-handed maximum control forces will be applicable. In other cases (such as takeoff rotation, or maneuvering during en route flight), the two-handed maximum forces will apply.
(2) Short-term and long-term forces should be interpreted as follows:
(a) Short-term forces are the initial stabilized control forces that result from maintaining the intended flight path following configuration changes and normal transitions from one flight condition to another, or from regaining control following a failure. It is assumed that the pilot will take immediate action to reduce or eliminate such forces by re-trimming or changing configuration or flight conditions, and consequently short-term forces are not considered to exist for any significant duration. They do not include transient force peaks that may occur during the configuration change, change of flight conditions, or recovery of control following a failure.
(b) Long-term forces are those control forces that result from normal or failure conditions that cannot readily be trimmed out or eliminated.
(3) In conducting the controllability and maneuverability tests of §§ 25.143, 25.145 and 25.147, consideration should be given to the requirements of § 25.671(c). Section 25.671(c) requires the airplane to “be shown by analysis, tests, or both, to be capable of continued safe flight and landing after any....failures or jamming in the flight control systems and surfaces (including trim, lift, drag, and feel systems), within the normal flight envelope....” This would include any single failure, or any combination of failures not shown to be extremely improbable.
The specific concerns for possible flight test evaluation are as follows:
(a) Unless safe flight characteristics are demonstrated with functionally related high lift or drag devices retracted on one side of the airplane and extended on the other, the
82 motion of the devices on opposite sides of the plane of symmetry must be synchronized by a mechanical interconnection or approved equivalent means as required by § 25.701 (Flap and slat interconnection).
(b) The “safe flight characteristics” referred to in paragraph (a), above, are interpreted as being the retention of adequate stall margins and controllability, following the described failure conditions, without requiring exceptional piloting skill or strength.
(c) Additionally, no hazardous change in altitude, or attitude, should occur during transition to the asymmetric condition considering likely transition rates.
(d) Cases of primary flight control jams (i.e., pitch, roll, and yaw) to be considered should include the following:
1 For pitch controls:
(aa) A jam at the initiation of the takeoff rotation, and
(bb) A jam at the most critical peak elevator angle reached during a normal rotation maneuver.
2 For roll and yaw controls, considered separately, a jam in the control position necessary to initiate all normal turns.
(4) Modern wing designs can exhibit a significant reduction in maximum lift capability with increasing Mach number. The magnitude of this Mach number effect depends on the design characteristics of the particular wing. For wing designs with a large Mach number effect, the maximum bank angle that can be achieved while retaining an acceptable stall margin can be significantly reduced. Because the effect of Mach number can be significant, and because it can also vary greatly for different wing designs, the multiplying factors applied to VSR may be insufficient to ensure that adequate maneuvering capability exists at the minimum operating speeds. To address this issue, § 25.143(h) was added by amendment 25-108 to require a minimum bank angle capability in a coordinated turn without encountering stall warning or any other characteristic (including the envelope protection features of fly-by-wire flight control systems or automatic thrust increases) that might interfere with normal maneuvering. The maneuvering requirements consist of the minimum bank angle capability the FAA deems adequate for the specified regimes of flight combined with additional bank angle capability to provide a safety margin for various operational factors. These operational factors include both potential environmental conditions (e.g., turbulence, wind gusts) and an allowance for piloting imprecision (e.g., inadvertent overshoots). The FAA considers the automatic application of thrust by an envelope protection feature to be a feature that might interfere with normal maneuvering because it will result in a speed increase and a flight path deviation, as well as potentially increasing crew workload due to the unexpected thrust increase.
b. General Test Requirements .
(1) Compliance with § 25.143 (a) through (g) is primarily a qualitative determination by the pilot during the course of the flight test program. The control forces required and airplane response should be evaluated during changes from one flight condition to another and during maneuvering flight. The forces required should be appropriate to the flight condition being evaluated. For example, during an approach for landing, the forces should be light and the airplane responsive in order that adjustments in the flight path can be accomplished with a minimum of workload. In cruise flight, forces and airplane response should be such that inadvertent control input does not result in exceeding limits or in undesirable maneuvers.
Longitudinal control forces should be evaluated during accelerated flight to ensure a positive stick force with increasing normal acceleration. Forces should be heavy enough at the limit load factor to prevent inadvertent excursions beyond the design limit. Sudden engine failures should be investigated during any flight condition or in any configuration considered critical, if not covered by another section of part 25. Control forces considered excessive should be measured to verify compliance with the maximum control force limits specified in § 25.143(d). Allowance should be made for delays in the initiation of recovery action appropriate to the situation.
(2) Since § 25.143(h) involves a target speed, bank angle, and maximum value of thrust, not all flight test conditions to demonstrate compliance will necessarily result in a constant-altitude, thrust-limited turn. In cases with positive excess thrust, a climbing condition at the target bank and speed is acceptable. Alternately, if desired, the thrust may be reduced to less than the maximum allowed, so that compliance is shown with a completely stabilized, constant-altitude turn. With the airplane stabilized in a coordinated turn, holding thrust and speed, increase bank angle at constant airspeed until compliance is shown. For cases with negative excess thrust (e.g., the landing configuration case), a constant-altitude slow-down maneuver at the target bank angle has been shown to be a suitable technique. With the airplane descending at VREF in wings-level flight on a three degree glide path, trim and throttle position are noted. The airplane is then accelerated to VREF + 10-20 knots in level flight. The original trim and throttle conditions are reset as the airplane is rolled into a constant-altitude slow-down turn at the target bank angle. Throttles can be manipulated between idle and the marked position to vary slow-down rate as desired. Compliance is shown when the airplane decelerates through VREF in the turn without encountering a stall warning or other characteristic that might interfere with normal maneuvering.
c. Controllability Following Engine Failure . Section 25.143(b)(1) requires the airplane to be controllable following the sudden failure of the critical engine. To show compliance with this requirement, the demonstrations described in paragraphs (1) and (2), below, should be made with engine failure (simulated by fuel cuts) occurring during straight, wings level flight. To allow for likely in-service delays in initiating recovery action, no action should be taken to recover control for two seconds following pilot recognition of engine failure. The recovery action should not necessitate movement of the engine, propeller, or trim controls, and should not result in excessive control forces. Additionally, the airplane will be considered to have reached an unacceptable attitude if the bank angle exceeds 45 degrees during the recovery. These tests may be conducted using throttle slams to idle, with actual fuel cuts repeated only for those tests found to be critical.
(1) At each takeoff flap setting at the initial all-engine climb speed (i.e., V 2 + 10 knots) with:
(a) All engines operating at maximum takeoff power or thrust prior to failure of the critical engine;
(b) All propeller controls (if applicable) in the takeoff position;
(c) The landing gear retracted; and
(d) The airplane trimmed at the prescribed initial flight condition.
(2) With the wing flaps retracted at a speed of 1.23 V SR with:
(a) All engines operating at maximum continuous power or thrust prior to failure of the critical engine;
(b) All propeller controls in the en route position;
(c) The landing gear retracted; and
(d) The airplane trimmed at the prescribed initial flight condition.
d. Airplane-Pilot Coupling (A-PC) .
(1) Explanation .
(a) Section 25.143(a) and (b) require that the airplane be safely controllable and maneuverable without exceptional piloting skill and without danger of exceeding the airplane limiting load factor under any probable operating conditions. Service history events have indicated that modern transport category airplanes can be susceptible to airplane-pilot coupling under certain operating conditions and would not meet the intent of this requirement.
(b) The classic A-PC situation, commonly referred to as a “pilot-induced oscillation” (PIO), is considered to occur when an airplane’s response is approximately 180 degrees out of phase with the pilot’s control input. However, A-PC events with 180 degrees phase relationships are not the only conditions in which the airplane may exhibit closed-loop (pilot-in-the-loop) characteristics that are unacceptable for operation within the normal, operational, or limit flight envelopes. Others include unpredictability of the airplane’s response to the pilot’s control input. This may be due to nonlinearities in the control system, actuator rate or position limiting not sensed by the pilot through the flight controls, or changing pitch response at high altitude as the airplane maneuvers into and out of Mach buffet. Artificial trim and feel systems which produce controllers with too small a displacement and light force gradients may also lead to severe over control. This is especially true in a dynamic environment of high altitude turbulence or upsets in which the autopilot disconnects. This places the airplane in the hands of the unsuspecting pilot in conditions of only a small g or airspeed margin to buffet
85 onset and with very low aerodynamic damping. These characteristics, while not 180 out of phase PIOs per se, may be hazardous and should be considered under the more general description of airplane-pilot coupling tendencies
(c) Some of the A-PC tendency characteristics described in paragraph (b) above are attributes of transport airplanes (e.g., low frequency short period, large response lags) that are recognized by part 25. Limits are placed on some of these individual attributes by part 25
(e.g., stick force per g, heavily damped short period) to assure satisfactory open-loop characteristics. However, service reports from recent years have indicated that certain operating envelope conditions, combined with triggering events, can result in airplane-pilot coupling incidents. Some of the conditions that have led to these A-PCs include fuel management systems that permit extended operations with a center of gravity at or near the aft limit, operating at weight/speed/altitude conditions that result in reduced margins to buffet onset combined with tracking tasks such as not exceeding speed limitations and severe buffet due to load factor following an upset, and control surface rate or position limiting.
(d) This service experience has shown that compliance with only the quantitative, open-loop (pilot-out-of-the loop) requirements does not guarantee that the required levels of flying qualities are achieved. Therefore, in order to ensure that the airplane has achieved the flying qualities required by § 25.143(a) and (b), the airplane should be evaluated by test pilots conducting high-gain (wide-bandwidth), closed-loop tasks to determine that the potential of encountering adverse A-PC tendencies is minimal.
(e) For the most part, these tasks should be performed in actual flight. However, for conditions that are considered too dangerous to attempt in actual flight (i.e., certain flight conditions outside of the operational flight envelope, flight in severe atmospheric disturbances, flight with certain failure states, etc.), the closed loop evaluation tasks may be performed using a motion base high fidelity simulator if it can be validated for the flight conditions of interest.
(2) Special Considerations .
(a) The certification team should understand the flight control system and airplane design.
(b) The applicant should explain why the design is not conducive to an A-PC problem and how this is to be shown in both developmental and certification flight tests.
(c) The applicant should explain what has been done during the development flight test experience and any design changes that were required for A-PC problems.
(d) The certification flight test program should be tailored to the specific airplane design and to evaluate the airplane in conditions that were found to be critical during its development program and A-PC analytical assessment.
(e) The FAA flight test pilots should also continuously evaluate the airplane for A-PC tendencies during the certification program in both the airplane and simulator. This
86 evaluation should include both normal and malfunction states; all certification flight test points;
transitions between and recoveries from these flight test points; and normal, crosswind, and offset landing task evaluations.
(f) Since the evaluation of flying qualities under § 25.143(a) and (b) is basically qualitative, especially evaluations of A-PC susceptibility, the high-gain tasks discussed herein should be accomplished by at least three test pilots. Use of other pilots can provide additional insights into the airplane handling qualities, but for the purpose of demonstrating compliance with this requirement the evaluation pilots should be trained test pilots.
(3) Procedures (Flight Test).
(a) Evaluation of the actual task performance achieved, e.g., flight technical error, is not recommended as a measure of proof of compliance. Only the pilot’s rating of the A-PC characteristics is needed as described in paragraph 20d(6). The tasks are used only to increase the pilot’s gain, which is a prerequisite for exposing A-PC tendencies. Although task performance is not used as proof of compliance, task performance should be recorded and analyzed to insure that all pilots seem to be attempting to achieve the same level of performance.
(b) Tasks for a specific certification project should be based on operational situations, flight testing maneuvers, or service difficulties that have produced A-PC events. Task requirements for a specific project will be dictated by the particular airplane and its specific areas of interest as determined by the tailored flight test program mentioned above. Some of these include high altitude upset maneuvers, encounters with turbulence at high altitude in which the autopilot disconnects, crosswind/crossed control landings with and without one engine inoperative, and offset landings to simulate the operational case in which the airplane breaks out of instrument meteorological conditions (IMC) offset from the glideslope and/or localizer beam and the pilot makes a rapid alignment correction. Tests should be conducted at or near the critical altitude/weight/c.g. combinations.
(c) Tasks described here may be useful in any given evaluation and have proven to be operationally significant in the past. It is not intended that these are the only tasks that may be used or may be required depending on the scope and focus of the individual evaluation being conducted. Other tasks may be developed and used as appropriate. For example, some manufacturers have used formation tracking tasks successfully in the investigation of these tendencies. For all selected tasks, a build-up approach should be used and all end points should be approached with caution. Capture tasks and fine tracking tasks share many common characteristics but serve to highlight different aspects of any A-PC problem areas that may exist.
In some cases, depending on individual airplane characteristics, it may be prudent to look at capture tasks first and then proceed to fine tracking tasks or combined gross acquisition (capture) and fine tracking tasks as appropriate.
(4) Capture Tasks .
(a) Capture tasks are intended to evaluate handling qualities for gross acquisition as opposed to continuous tracking. A wide variety of captures can be done provided the