Document
F
AIAA-92-4107-CP
PITCH CONTROL MARGIN AT HIGH ANGLE OF ATT ACK - QUANTITATIVE REQUIREMENTS (FLIGHT TEST CORRELATION WITH SIMULATION PREDICTIONS) J. Lackey Capt. C. Hadfield (CAF) Naval Air Warfare Center - Aircraft Division Patuxent River, Maryland Q (at 2 sec) pitch rate at two seconds from recovery input ABSTRACT
qoor
dynamic pressure pitch acceleration qoot Aircraft designs that employ relaxed static stability Qdot (at l sec) pitch acceleration at one second from (RSS) have the following problem: reduced pitching recovery input moments associated with RSS at high angle of attack Qdot (avg:<:::;l sec) average pitch acceleration (AOA) require a minimum pitch recovery moment or within one second from recovery margin to guarantee a safe return from high AOA input maneuvers at the most aft center of gravity (CG) Qdot (max:<::;! sec) maximum pitch acceleration within encountered during a mission. Recent incidents and one second from recovery input mishaps on Class IV aircraft have demonstrated a need pitch rate change within a time for establishing quantitative longitudinal high AOA pitch control margin design guidelines for future interval
aircraft. The Naval Air Warfare Center - Aircraft s reference wing area
Division (NAWC-AD) is currently supporting an effort Tree time to recover to less than 10 deg in conjunction with NASA Langley Research Center AOA (NASA LaRC) to quantify such requirements. NASA a angle of attack LaRC has conducted a series of extensive simulation y flight path angle rate evaluations to define these design guidelines. The .1AOAAt angle of attack change within a time purpose of flight tests were to validate the overall interval research test methodology by comparing pilot altitude required to recover .1hrec comments, pilot ratings, and aircraft response airspeed change within a time .1V At characteristics gathered during inflight recoveries from interval high AOA conditions to those gathered during the fixed- pitch angle change within a time .1E> At base simulation sessions. Tests were completed on an F/A-18A in six flights for a total of 9.8 flight hours interval
e pitch angle
using an AOA and CG buildup sequence. Flight test results have validated the simulation studies in that pilot rating of high AOA nose-down recoveries were based on the short-term response interval in the forms of pitch acceleration and rate. In addition, flight test has INTRODUCTION demonstrated that high AOA pitch control margin can be evaluated using a stabilized pushover method.
In the Navy, there is currently an effort underway in conjunction with NASA Langley Research Center NOMENCLATURE (NASA LaRC) to define quantitative longitudinal high AOA pitch control power/ margin requirements. Initial AOA angle of attack work to define such guidelines was conducted from AOAdotAt angle of attack rate change within a November 1989 to June 1990 at NASA LaRC by a Navy / NASA LaRC team using both a baseline and time interval modified parametric F/A-18A six degree of freedom cbar mean aerodynamic chord simulation model in the fixed-base Differential CG center of gravity Maneuvering Simulator (DMS). A Pitch Recovery Cm total pitching moment Rating (PRR) scale (see figure I) was developed to ftHp pressure altitude feet correlate qualitative pilot opinion with nose-down pitch Iyy pitch inertia response characteristics of an aircraft. Navy / NASA KCAS knots calibrated airspeed LaRC simulation studies produced specific candidate MAC mean aerodynamic chord figures of merit to quantify high AOA longitudinal Release C: This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States pitch control margin requirements. In order to validate ScOJ)e of Tests simulation results, flight tests were planned for two phases. Phase I tests, conducted from 30 September to 8 Simulation tests were conducted on the NASA October 1991, consisted of a limited study using an LaRC DMS using a total of six pilots for 55 test Fl A-18 to validate the overall research test hours. Out of the six pilots, two pilots conducted the methodology. Phase II tests will consist of a more Phase I flight tests (designated Pilot A and Pilot B) detailed approach emphasizing guideline validation completing a total of 12 and 8 simulation test hours, using the NASA Dryden F/A-18 High Alpha Research respectively. Simulation tests were conducted in two Vehicle in which flight test flight control laws can be phases. Phase A tests consisted of evaluation modified as desired in conjunction with thrust vectoring methodology development using a baseline F/A-18 controls. The Naval Air Systems Command simulation model to vary nosedown response with CG (NA VAIRSYSCOM) tasked NA WC-AD to conduct the movement. Phase B tests consisted of developing Phase I tests. This paper outlines the Phase I test candidate guidelines via parametric study in which results.
variation of selected pitching moment parametrics allowed evaluation pilots to rate high AOA recoveries at Description of Test Aircraft more varied response conditions. The parametric studies were conducted on a modified Ff A-18 simulation model, details of which arc presented in reference 2.
The F/A-18A (see figure 2) is a single seat, high performance, twin engine supersonic fighter Preflight ground tests were conducted at the characterized by moderately swept, variable camber mid- NA WC-AD Aircraft Test and Evaluation Facility mounted wings, twin outboard canted vertical stabilizers (A TEF) to determine the empty weight and moment mounted forward of the horizontal stabilators, a values for the test loading, to calculate CG error at full, spccdbrake located on the upper aft section of the half-full, and empty fuel states by comparing true fuselage between the vertical stabilizers, and leading values calculated at ATEF with values determined via edge extensions mounted on each side of the fuselage telemetry readings of individual fuel tank quantities, and from the wing roots to just forward of the windshield.
to ensure that the nonproduction CG control system The airplane is configured with full span leading edge flaps, inboard trailing edge flaps, and outboard ailerons worked properly.
on each wing. The flight control system consists of two digital flight control computers that utilize a full A total of 6 flights for 9.8 flight hours were completed by two evaluation pilots during this authority control augmentation system to operate the evaluation. The flights were conducted in two phases.
hydraulically driven control surfaces. The test airplane Phase I A tests were flown to: (1) ensure that the test was equipped with version 8.3.3 programmable read airplane was rigged properly to minimize roll-off only memory flight control laws. The aircraft is powered by two General Electric F404-GE-400 tendencies at high AOA, (2) allow the pilots to become familiar with the test maneuver at forward CG positions augmented turbofan engines rated at 16,000 pounds (:<,; 23% MAC) through an AOA buildup range, and (3) maximum uninstalled static sea level thrust. A detailed practice using the CG control system. All phase IA description of the F/ A- l 8A airplane is presented in tests were conducted within reference 2 limits. Phase 1B reference 1.
consisted of tests that varied the magnitude of pitch control margin available at target AOA's of 40 and 50 Description of Test Eguipment and Instrumentation deg using various CG positions (22.5 - 26.5 %MAC).
Phase IB tests were conducted outside of reference 2 A Nose Instrumentation Pallet System was limits as authorized by an approved NA V AIRSYSCOM installed in the airplane in order to transmit selected flight clearance. All tests were conducted in the cruise 1553 multiplex bus parameters to the real-time configuration as defined by gear up, flaps AUTO, telemetry processing system (RTPS) for monitoring speedbrakc retracted, and thrust as required to maintain during the tests. AOA was obtained from both the test conditions. All tests were conducted in the clean production air data computer and the inertial navigation loading as defined by no stores or pylons on any loading system (INS). Angle of sideslip was obtained from the stations.
INS. A non-production CG control system wa<; installed so that CG position could be changed by the pilot by Method of Tests selectively disabling fuel transfer using motive flow shutoff valves from a cockpit mounted control panel All test maneuvers consisted of symmetrical, (sec figure 3). The shutoff valves controlled fuel transfer stabilized lg trim pushovers and were conducted from from the forward and aft fuel tanks into engine feed various AOA's and CG's at an initial pitch attitude of tanks (see figure 4). The test airplane was not equipped 15 degrees (sec figure 5). The tests were not conducted with a flight test noseboom or a nonproduction backup "blind" (i.e. pilot knew aircraft CG position for safety emergency system (i.e. spin recovery chute).
of flight purposes). Test maneuvers were flown in the NA WC-AD local North and South "Spin" areas during daylight visual meteorological conditions. All flights were flown with a safety chase. Telemetry data were
T
transmitted via pulse code modulation received at 20 pitch rate was found to best correlate with pilot rating in the form of pitch rate at two seconds from recovery samples per second to RTPS where NA WC-AD engineers directed and monitored the tests. input (Q (at 2 sec)). The simulation studies defined time to recover as the time to reduce AOA to less than IO RES ULTS AND DISCUSSION degrees (Tree) because for typical tactical aircraft, this marks the central region of the low AOA operational envelope.
Figures of Merit Maximum Pitch Acceleration within one from The development of quantitative nose-down pitch control margin guidelines required the establishment of Recovery Input (Odot {max < I sec}) figures of merit to be used in evaluating recovery characteristics. A large number of candidate figures of Flight test matched simulation well only at lower merit were considered during the Navy / NASA LaRC pilot ratings (~ 3) (see figures 7 and 8). Higher pilot simulation studies. The key to establishing their ratings exhibited significant flight test to simulation importance with respect to control margin (power) was divergence. These differences can be explained by pitch to chronologically order the parameters relative to acceleration nonlinearities produced due to the flight initiation of recovery controls (see figure 6). In figure 6, control system, aerodynamic effects, and motion cue as one progresses from short to long on the time scale, effects. The observed flight test pitch acceleration the figure of merit correlation with control margin nonlinearities explain differences between parametric decreases. Thus, those figures of merit on the left side simulation results because as a result of modifying the of the scale would be expected to be more important for simulation as presented in reference 2, flight control nose-down control design. During the simulation logic and modelled aerodynamics were fixed such that studies it was found that the angle of attack figures of nosedown recoveries exhibited "ideal" (no reversal) merit (AOAdotM and ~AOA~t) and the pitch attitude linear pitch acceleration responses. When rating the flight test maneuvers, evaluation pilots observed the figure of merit, ~Q~t, were poor correlators because the nonlinear tendencies as undesirable rate hesitation, evaluation pilots tended to rely more on out-of-the- producing higher pilot ratings as a result. Differences cockpit, visual cues (i.e. pitch accelerations and rates) between the baseline F/A-18 simulation and flight test during the recoveries vice looking for changes in AOA are primarily due to motion cue effects. Pilot comments and pitch attitude readings within a certain period of indicated motion cues were very important when time. The two figures of merit, ~ V ~t and ~hrec, were assessing immediate pitch response inflight. Through also found to be poor correlators because they tended to motion cues, degraded pitch responses were more be based more on airframe performance than control evident and made the evaluation pilots more critical of power. The remaining figures of merit, qdot, q~t and desired response than in the simulator where nosedown Tree were subsequently chosen as candidate figures of response cockpit cues were limited to the HUD and merit for the phase I flight tests.
dome visuals. It should be noted that the evaluation pilots knew aircraft CG due to safety of flight purposes During the NASA LaRC simulation studies, it was and thus had an idea of upcoming aircraft nose-down determined by the evaluation pilots that pitch response tendency. In conclusion, Qdot (max ~I sec) acceleration was the most strongly perceived nose-down was found to have good correlation between simulation response cue. In the absence of significant angular and flight test at lower pilot ratings (~ 3) where flight rates, pitch acceleration is strongly related to an test maneuvers exhibited "ideal" simulation cases via aircraft's pitch control power due to the direct linear pitch acceleration response. Increasing aft CG proportionality to static pitching moment (equation resulted in increasing pitch acceleration nonlinearities (I)), which contributed to flight test data divergence from predicated simulation pilot rating trends. Motion cue qdot = Cm * (qbar*S*cbar) effects became apparent at the higher pilot ratings where Iyy (I) increased pilot sensitivity to degraded pitch responses resulted in more critical ratings than compared to Since pitch acceleration was one of the first parameters simulation.
perceived by the pilots during a pushover recovery from high AOA (within the first second of the recovery), it Pitch Rate At Two Seconds From Recovery Input was considered as the most important figure of merit (0 (at 2 sec)) when attempting to quantify longitudinal control margin requirements. During the simulation studies Variations of pitch rate with pilot rating and CG pitch acceleration was found to correlate best with pilot position are shown in figures 9 and 10. Pilot A nose- rating in the form of maximum pitch acceleration down pitch rates tended to be higher than Pilot B values within one second from recovery input (Qdot (max at essentially the same CG's (gross weights) because of ~lsec)). Pilot comments also indicated that in addition differences in dynamic pressure where stabilized to initial pitch acceleration, pitch rate around two pushovers were conducted at lower altitudes. Flight test seconds from recovery input was also used in the pitch Q (at 2 sec) agrees with predicted simulation pilot rating recovery rating process. Simulation results showed that trends only at lower pilot ratings (~ 3). A more aft CG
l
decreases the static pitching moment. However in the Pitch Acceleration At One Second From Recovery F/A-18, pitch rate response remains essentially InIJul {Odot {at I sec)) invariant due to the effects of AOA and pitch rate feedback in the flight control system. This flight Variation of Qdot (at I sec) with pilot rating is control system effect can be observed in figure 11 in presented in figures 15 and 16. Flight test Qdot (at 1 which for the full forward stick recoveries, stabilator sec) values tended to match the simulation better at saturation duration varies with CG. In conclusion, Q (at higher pilot ratings for each evaluation pilot than those 2 sec) was found to have good correlation between observed with Qdot (max $I sec). This improved match simulation and flight test at lower pilot ratings (s 3).
supports the argument that the evaluation pilots referred The fact that flight test values of Q (at 2 sec) were to the pitch acceleration nonlinearities when essentially constant for pilot ratings from 2 to 4.5 determining final ratings. However, the degree of Qdot indicates that (I) pitch rate effects were secondary in (at I sec) data scatter per pilot rating was larger than determining overall response rating and / or (2) pitch those observed with qdot (max $1 sec). The fact that rate in the form of Q (at 2 sec) is not the best there exists a larger amount of vertical data scatter (both correlating case.
in the case of simulation and flight test) indicates that evaluation pilots did not rate the response by solely Time To Recover {Tree) using qdot (at 1 sec) in their overall assessments and this figure of merit is not very consistent. In The variations of Tree with pilot rating are shown conclusion, flight test Qdot (at I sec) values were found in figures I 2 and 13. Pilot comments indicated that Tree to correlate better with simulation at higher pilot was never strongly perceived during the pushovers.
ratings; however, excessive vertical data scatter per When comparing flight test results to simulation data, rating indicates that this figure of merit is not very Tree matched fairly well; in both cases it was strong.
characterized by essentially negligible variations with pilot rating except in extremely degraded response (high AOA hangup-type) cases which were only investigated Average Pitch Acceleration Within One Second during simulation for safety of flight purposes.
From Recovery Input (Odot (avg <l sec)) Essentially constant Tree up until the very high pilot ratings (4.5 to 5) indicates that it is more long term, Variation in Qdot (avg $1 sec) with pilot rating is hang-up response related, in contrast to pitch shown in figures 17 and 18. Flight test Qdot (avg $1 acceleration and rate which are short term, normal sec) and simulation values exhibited considerably recovery related. In conclusion, Tree was found to have reduced vertical data scatter; however, the magnitude of good correlation between simulation and flight test in the gradient with respect to pilot rating is small and that minimal variation of this figure of merit was some flight test to simulation data divergence is observed during flight tests at low to high pilot ratings apparent at the higher pilot ratings(~ 4). Differences in (2 to 4).
data can be explained by considering that the previously discussed pitch acceleration nonlinearities are being Other Pitch Acceleration Figures Of Merit averaged into this figure of merit. In conclusion, Qdot (avg $1 sec) was found to have low vertical data scatter per pilot rating; however, it is a poor figure of merit General considering that overall variation with pilot rating was small.
Figures 7 and 8 indicate small variation of flight test Qdot (max $I sec) with pilot rating. However, Workload Required For Test Maneuver Stabilization figure 14 shows significant variation of pilot rating with CG. This clearly indicates that the pitch The maneuver test method required that evaluation acceleration figure of merit, Qdot (max $!sec), is pilots vary thrust to stabilize at a constant pitch somewhat weak in not accounting for the nonlinear attitude. During flight test it was found that responses as discussed previously. The pilot is establishing required test conditions using this method obviously seeing degradation in pitch response, but this was very difficult since the pilot had to "close-the-loop" effect is not being reflected by Qdot (max $I sec). It on trim airspeed with throttles to keep flight path angle should be emphasized that during flight tests, the rate zero. Pilot B commented that "airspeed control evaluation pilots knew aircraft CG and thus were better through throttle adjustments was difficult due to large able to predict stabilized pushover response trends. In an -1 S KCAS airspeed jumps and strong airspeed effort to correlate data more closely, two other pitch sensitivity to thrust"; he further stated that "this effort acceleration figure of merit forms were investigated.
distracted allcntion from the initial portion of the One form was pitch acceleration al one second from pushover and may have affected pilot ratings. In recovery input (Qdot (at I sec)) and the other was simulation, entry conditions were automatic and average pitch acceleration within one second from effortless so all attention was focused on the pushover".
recovery input (Qdot (avg sl sec)).
Since the NASA LaRC simulation studies, as documented in reference 2, indicated that initial pitch delineating the definition of an undoubtful versus attitude had a negligible effect on pilot ratings, an alternative approach would be to hold constant thrust doubtful recovery.
and vary pitch attitude to stabilize. Using this approach, the maneuver set-up may be easier with reduced pilot workload and could result in more repeatable results.
During flight tests, a stabilized pushover was conducted CONCLUDING REMARKS in this manner to compare pilot workload. When approaching a stabilized pushover at constant pitch Results of the Navy / NASA LaRC pitch control attitude, significant pilot workload was evident from margin simulation studies were validated in that pilot continuous throttle inputs made while decelerating from cueing (rating) of high AOA nose-down recoveries 35 deg AOA to the target 50 deg AOA condition. As during flight test was based on the short-term response AOA continuously increased during the deceleration, in the forms of pitch acceleration and rate figures of increased thrust inputs were required to maintain zero merit. The final figures of merit forms to quantify high flight path angle rate. In addition to the various throttle AOA pitch control margin requirements, however, are inputs, continuous longitudinal stick inputs were yet to be determined. Flight test proved that high AOA required to hold the target pitch attitude at 15 deg. When pitch control margin can be demonstrated using a conducting the maneuver at a constant thrust setting stabilized pushover method; however, improvements in (MIL power), however, pilot workload was reduced from method technique are warranted. Once modifications are two (longitudinal stick and throttles) to one completed, the pitch recovery rating scale will be a vital (longitudinal stick) input controllers. During the start of tool in quantifying desired pitch control margin during the deceleration, the pilot simply pulled to 35 deg pitch future simulation and follow-on flight test evaluations.
attitude, set thrust to MIL and progressively pulled aft stick to increase AOA, sacrificing pitch attitude in the Specific conclusions established during the tests were as process of maintaining a stabilized condition with zero follows: flight path angle rate.
a. Qdot (max $1 sec) was found to have good Pitch Recovery Rating Scale Improvements correlation between simulation and flight test at lower pilot ratings ($ 3) where flight test maneuvers exhibited Flight tests showed that improvements in the pitch "ideal" simulation cases via linear pitch acceleration recovery rating scale used during the figure of merit response. Increasing aft CG resulted in increasing pitch correlations should be considered. The first weak area acceleration nonlinearities which contributed to flight found was mission task ambiguity. Using the PRR test data divergence from predicated simulation pilot scale, as defined from the simulation studies, required rating trends. Motion cue effect'> became apparent at the that each evaluation pilot generate his own mission higher pilot ratings where increased pilot sensitivity to scenario to rate the quality of pushovers when degraded pitch responses resulted in more critical ratings recovering from high AOA. Pilot A used: "vertical compared to simulation.
fight, coming uphill offensively, realizing late that I don't have enough energy to make it over the top, and b. Q (at 2 sec) was found to have good correlation unloading with full forward stick to gain energy as a between simulation and flight test at lower pilot ratings bogey moves into a position of advantage." Pilot B ($ 3). The fact that flight test values of Q (at 2 sec) used: "pushover from a nose high attitude to point were essentially constant for pilot ratings from 2 to 4.5 towards a bogey below". Pilots A and B were clearly indicates that (1) pitch rate effects were secondary in rating the maneuvers from different mission viewpoints. determining overall response rating and / or (2) pitch A more objective scenario should be used in which pilot rate in the form of Q (at 2 sec) is not the best ratings use a more standardized mission environment.
correlating case.
Another weak area found by the evaluation pilots was that the decision trees used in the PRR scale were too c. Tree was found to have good correlation between ambiguous. The reference 2 proposed PRR scale simulation and flight test in that minimal variation of revisions (shown in figure 19) more clearly define the this figure of merit was observed during flight tests at decision factors involved, particularly with respect to low to high pilot ratings (2 to 4).
adequacy of safety and a tactically desirable response.
Since the prime area of interest in the PRR scale when d. Qdot (at 1 sec) flight test values were found to establishing specification requirements is in the 4 to 5 correlate better with simulation at higher pilot ratings; rating region, this area needs to be expanded to more however, excessive vertical data scatter per rating clearly define the boundary between an undoubtful and indicates that this figure of merit is not very consistent.
doubtful recovery. The reference 3 proposed PRR scale revisions (shown in figure 20) expand this critical area. e. Qdot (avg :<::;l sec) was found to have low vertical In conclusion, the PRR scale as defined from the initial data scatter per pilot rating; however, it is a poor figure simulation studies was found to be weak in not defining of merit considering that overall variation with pilot a standardized mission scenario, using ambiguous rating was small.
decision tree factors to obtain ratings, and not clearly f. The PRR scale as defined from the initial simulation studies was found to be weak in not defining a standardized mission scenario, using ambiguous decision tree factors to obtain ratings, and not clearly delineating the definition of an undoubtful versus doubtful recovery.
Specific recommendations established during the tests were as follows: a. Recommended that further flight tests be conducted using a constant thrust, varying pitch attitude stabilized pushover technique to determine the degree of difficulty to conduct such a maneuver and define the effects of varying initial pitch attitude on pilot ratings.
b. Recommend that a standardized mission scenario be adopted for the PRR scale, the scale be restructured with the reference 2 decision tree revision recommendations, and that the scale be expanded in the 4 to 5 rating region per the reference 3 proposal.
c. Recommend further simulation and flight test studies be conducted to determine a pitch acceleration figure of merit which will better account for a wide range of pitch response conditions in a consistent manner.
REFERENCES 1. NATOPS Flight Manual Navy Model F/A- 18A/B/C/D, Al-Fl8AC-NFM-OOO, 15 January 1991.
2. NA V AIRW ARCENAIRDIV Tech Memo, SA- 167R-91, McNamara W.G., et al; Navy High Angle Of Attack Pitch Control Margin Requirements for Class IV Aircraft, April 1992.
3. Hadfield, C., High Angle of Attack Handling Qualities Rating Scales, Thesis presentation for Master of Science degree from University of Tennessee, Knoxville Tennessee, May 1992.
J
Enhancing for mission Sausfactory for m1ss1on Degrades mission sltghlly Moderate! y degrades m ISSIOO and I or safety Decision Facton Notes: 1. W ~ there enougn pncn response 7 (I) Pilot Rating= 2.5 (Min. Tactically Desirable) Level A 2. Could you use more response?
(2) Pilot Rating= 4.5 (Min. Safety Level) Level B Pilot Decisions I ◄--- 3. Wu time to recover short enough?
._ ______ _. 4. Was the recovery in question?
5. Wu pilot compenution required?
6. ls the response suitable for the mission?
Figure 1 - Pitch Reco~ery R~ting Sc~e
(Developed Prior To S1mulat1on Studies) Figure 2- F/A-18 Hornet
s
s
MANUAL CG CONTROL MANUAL FWD .------, I I I
~I
I
(j)
! I I I ______ J AFf NORMAL PITCH CONTROL MARGIN
s
s
Figure 3 - CG Control System Cockpit Control Panel Forward Tank4 Tank3 Tank2 Tankl From left and right From left and right internal wing tanks internal wing tanks To right engine pump To left and right To left engine internal wing tanks Nole: - Transfer motive flow
1r Transfer ejector pump
Not shown: I. Hot fuel recirculation - Engine Feed 2. Vent/ dump system u,u.,,-,, Tank I and 4 transfer
8 Engine feed turbo pump
3. External fuel system .,. '- '- Engine feed motive flow 4. Refuel / defuel system c:::J Fuel line connector - lntemal wing transfer Figure 4 - F/A-18A Fuel System Schematic Velocity Figure 5 - Stabilized Pushover strong C
-~
~
g
§
u
..c ii; g Q.
q ~t
:.a
qdJt Cl) MOA C ~t
.g .:1V ~t
AOAdot ~t .:18 ti.t <':I ~ .:1hrec Tree 0::: weak short Time Scale long Figure 6 - Figures of Merit ◄ HJ Pilot A Flight Test 6 Pilot A Baseline Simulation A Pilot A Parametric Simulation ----B----- Preliminary Guideline (Simulation) -0.1 -0.4 5 6 3 4 PILOT RA TING Figure 7 - Qdot (max ~1 sec) versus Pilot Rating (Pilot A) X Pilot B Flight Test A Pilot B Parametric Simulation -e-- Preliminary Guideline (Simulation) -0.1 -0.2 VI N'
if
0~ -0.3 -0.4 -0.5 2 3 4 5 6 PILOT RA TING Figure 8 - Qdot (max ~1 sec) versus Pilot Rating (Pilot B) EE Pilot A Flight Test b. Pilot A Baseline Simulation .6. Pilot A Parametric Simulation ~ Preliminary Guideline (Simulation) 2 3 4 5 6 PILOT RA TING
Figure 9 - Q (at 2 sec) versus Pilot Rating (Pilot A)
X Pilot B Flight Test .6. Pilot B Baseline Simulation ~ Preliminary Guideline (Simulation) 0.0 -10.0 -20.0
1~
c::!l
$-8 0~ -30.0 -40.0 -50.0 1 2 3 4 5 6 PILOT RA TING Figure 10 - Q (at 2 sec) versus Pilot Rating (Pilot B)
-·-•4---
EB Pilot A Right Test X Pilot B Fligh1 Test 3.0 2.5
~
z 2.0 {:: < L5
g~
~M, LO
°'
0.5 ;s < !- <n 0.0 2.2 23 24 25 26 27 CENTER OF GRAVITY (%MAC) Figure 11 - Stabilator Saturation Time versus CG Ell Pilot A Flighl Test 6 Pilot A Baseline Simulation .l Pilot A Parametric Simulation 25,0 20.0 15.0
&i~
f= !
10.0 5.0 0.0 3 4 PILOT RA TING Figure 12 - Tree versus Pilot Rating (Pilot A) X Pilot B Flight Test A Pilol B Baseline Simulation
-e-- Preliminary Guideline (Simulation)
25.0 20.0 15.0 u,..._ UJ ~
I= ~
10.0 5.0 0.0 2 4 5 3 6 PILOT RA TING Figure 13 - Tree versus Pilot Rating (Pilot B) B3 Pilot A Right Test X Pilot B Flight Test 5.0
! 4.0
~ f-····•·····'····
~ 3.0 1--···············
2.0 1.0 '---'---'-..!---'----'---'----'----'-~..k_..J._.J.._J.__1-....J-..1---L--'---'---'---'-_J._..k_-1-- 22 23 24 25 26 27 CEN'JER OF GRAVITY (%MAC) Figure 14 - Pilot Rating versus CG i a, EB Pilot A Flight Test t:,. Pilot A Baseline Simulation .t. Pilot A Parametric Simulation
-e-- Preliminary Guideline (Simulation)
0.0 -0.1
jR
-0.2
:: ~
-=-~
~~ -0.3 -0.4 -0.5 I 2 4 5 3 6 PILOT RATING Figure 15 - Qdot (at 1 sec) versus Pilot Rating (Pilot A) X Pilot B Flight Test .t. Pilot B Baseline Simulation
-e-- Preliminary Guideline (Simulation)
0.0 -0.1 -0.2
i~
ii
!~
-0.3 -0.4 -0.5 I 2 3 4 5 6 PILOT RATING Figure 16 - Qdot (at 1 sec) versus Pilot Rating (Pilot B)
p
m Pilot A Flight Test
.A Pilot A Baseline Simulation
--e-- Preliminary Guideline (Simulation)
0.0 -0.1 -0.2
~~
$~
15---- -0.3
B
-0.4 -0.5 2 3 5 6 PILOT RA TING Figure 17 - Qdot (avg $1 sec) versus Pilot Rating (Pilot A) X Pilot B Flight Test .A Pilot B Baseline Simulation -B-- Preliminary Guideline (Simulation) 0.0 -0.I
i,.._
-0.2 -<'I
if
cS ._, -0.3
B
-0.4 -0.5 1 2 4 5 3 6 PILOT RATING Figure 18 - Qdot (avg Sl sec) versus Pilot Rating (Pilot B)
l
◄
Enhancing for mission Satisfactory for miss,on was not a concern) Degrades mission slightly Adequale =""'Y Moderately degrades missioo and / or safety Decision Facton Not.es: l. Wu there cnougn pucn response?
(I) Pilot Rating= 2.5 (Min. Tactically Desirable) Level A 2. Coukt you use more response?
(2) Pilot Rating= 4.5 (Min. Safety Level) Level B Pilot Decisions 3. Was tune to recover short enough?
4. Wu the recovery in question?
5. Was pilot compensation required?
6. b the rcs]XmSC suitable for the mission?
Figure 19 - Revised Pitch Recovery Rating Scale (Developed After Simulation Tests Completed) DECISIO:- FA<.'TORS Wu there enough pitch rcsponK (accelerauon, rate)?
RA TING l£VEL5 2. Could you use more response?
P£RFORM TEST DEANE Lvl A. Tacl1caJly Desirable 3. Wu the tune to recover short enough?
MANEUVER MISSION SCENARIO Lvl B. Adequa~ for Safety 4. Wu recovery in question?
5. Wu pilot compcnsalion required?
fi. l!1 the re" n~c rn11ah\c for the mi~111on?
Figure 20 - Revised Pitch Recovery Rating Scale (Developed After Flight Tests Completed)