Piloted-simulation study of effects of vortex flaps on low-speed handling qualities of a Delta-wing airplane

- .

NASA

Technical

Paper

September 1987

Piloted- Simulation Study

of Effects of Vortex

Flaps - on Low-Speed

Handling- Qualities of

a Delta-king Airplane

Jay hl. Brandon,

Philip W. Bro~rxi, and

,4lfrcd J. b'unsche'l

NASA

Technical

Paper

Piloted-SimulationStudy

of Effects of Vortex

Flaps on Low-Speed

Handling Qualities of

a Delta-Wing Airplane

Jay M. Brandon and

Philip W. Brown

Langley Research Center Hampton, Virginia

Alfred J. Wunschel

U S . Air Force Systems Command

Langley Research Center Hampton, Virginia National Aeronautics and Space Administration Scientific and Technical I Information Office ~~~~ ~~~ ~ Summary Wind-tunnel tests to date have shown that the addition of vortex flaps to an airplane can signif- A piloted-simulation study was conducted to icantly impact its stability and control character- investigate the effects of vortex flaps on low-speed istics, and therefore flying qualities might also be handling qualities of a delta-wing airplane. The sim- affected. To investigate this effect, a piloted- ulation math model was developed from wind-tunnel simulation study was initiated to assess the flying tests of a 0.15-scale model of the F-106B airplane.

qualities of the F-106B with vortex flaps over the Pilot evaluations were conducted using a six-degree- planned test envelope. This report summarizes the of-freedom motion base simulator. The results of results of the simulation study for the critical low- the investigation showed that changes in lateral- speed approach-to-landing flight phase.

directional handling qualities due to the addition of vortex flaps were minimal; however, the reduced Symbols static longitudinal stability caused by the vortex All longitudinal data are referenced to the wind- flaps significantly degraded handling qualities in axis system, and all lateral-directional data are the approach-to-landing task. Acceptable handling I referenced to the body-axis system. All force data I qualities could be achieved by limiting the aft center- measurements were obtained at a center-of-gravity I of-gravity location, consequently reducing the oper- location of 27.5 percent of the wing mean aerody-

i ational envelope of the airplane. Further improve-

namic chord. Force coefficients for all configurations I ments were possible by modifying the flight control are based on the reference geometry of the basic force-feel system to reduce pitch-control sensitivity.

configuration.

Introduction b wing span, f t drag coefficient, Drag/ijS C D The vortex flap concept, illustrated in figure 1, lift coefficient, Lift/@ CL repositions the wing leading-edge vortices such that they move from the wing onto the flap surface. The rolling-moment coefficient, C L resulting induced suction pressures on the forward- Rolling moment/@% facing flap produce a thrust component that reduces roll-control power drag. This concept has been examined in many experimental and theoretical studies (refs. 1 to 4) and has been found to be an effective method for improving the performance of highly swept config-

= CiP + C i P . sin a, per radian

urations at maneuver lift coefficients. As a result, aC considerable research has been conducted to develop

= +, per degree

analytical tools for the design of efficient and highly effective vortex flaps (refs. 5 to 7).

2, per radian

A flight validation effort to verify the analyti- cal design tools and expected performance benefits pitching-moment coefficient, for the vortex flap concept was recently initiated at Pitching moment/@%

I

the NASA Langley Research Center. The airplane

chosen for the test was an F-106B. The reasons = a per radian

I a% , for selecting this airplane include (1) ease of struc- tural modifications to accommodate vortex flaps,

= Cmq + Cm,, per radian

(2) compatibility of the operating envelope of the

= %, per radian

F-106B with test requirements, and (3) compatibil- a, ity of the 6 0 ' delta wing with the vortex flap con- yawing-moment coefficient, cept. The flight validation effort is focused primar- Yawing moment/qSb ily on verification of the design procedures employed in the design of the flaps. This verification will be yaw-control power accomplished with comparisons of flight, wind- = +, aC per radian tunnel, and analytically predicted pressure distribu- a, tions over the wing. In addition, a limited amount of = Cn, - C cosa, per radian performance flight testing will be conducted to com-

na

pare the relative performance of the vortex flap and

= %, per degree

basic configurations.

= s, per radian

rudder deflection angle, deg damping ratio sideforce coefficient , phase angle of Dutch roll Side force/ijSb oscillation in sideslip

= %, per degree

W undamped natural frequency, rad/sec reference mean aerodynamic chord Subscripts: acceleration due to gravity DR Dutch roll ( l g % 32.17 ft/sec2) phugoid P height of E/4 above ground, f t short period SP moment of inertia about X-, Y-, and Z-axis, respectively, Abbreviations: slug-ft2 product of inertia, slug-ft2 I X Z c.g. center of gravity - - Roll command IFR instrument flight rules k Roll performance (flight with no outside visual references) lift-tedrag ratio LID roll rate about body X-axis, PR pilot rating P deglsec VFR visual flight rules (flight with outside visual references) pitch rate about body Y-axis, deg/= Airplane Description freestream dynamic pressure, A basic airplane configuration and vortex flap lb/ft2 configurations were compared to assess the effects of r yaw rate about Z-axis, deg/sec vortex flaps on handling qualities. The basic air- plane represented in the simulation study was an S reference wing area, ft2 F-lNB, which is a two-place, delta-wing, single engine jet fighter. A three-view sketch of the basic t time, sec configuration is shown in figure 2, and the mass and V free-stream velocity, ftlsec geometric characteristics used in the simulation are longitudinal, lateral, and listed in table I. The primary aerodynamic controls

x, y, z

include symmetric and antisymmetric deflection of vertical body axis, respectively the elevons for pitch and roll control and deflection cr angle of attack, deg of the rudder for yaw control. A speed brake was available to the pilot to modulate drag. The airplane

ir = $$$, radlsec

is equipped with an irreversible flight control system.

angle of sideslip, deg The control forces experienced by the pilot are pro- P vided by an artificial force-feel system that produces

= g, rad/=

P

pitch stick and rudder pedal forces that vary with dynamic pressure. The roll-control forces, however, maximum change in sideslip A&ax are invariant with flight condition. The forcefeel sys- occurring within 2 sec for a tem characteristics are shown in figure 3. The flight step roll-control command control system incorporates an elevon-rudd-- e l . coor- flight path angle, deg dination feature and a stability augmentation sys- tem (SAS) consisting of pitch and yaw rate dampers antisymmetric elevon deflec- (ref. 8). Normal pilot operating procedures, however, tion, deg call for deactivation of the SAS during the approach- symmetric elevon deflection, to-landing flight phase which is the primary focus of deg this investigation.

Wind-Tunnel Tests The vortex flap configurations were derived by the addition of constant-chord vortex flaps mounted Data were obtained for the simulation study from on the basic wing leading edge at deflection angles of the results of wind-tunnel tests on two models in 30°, 40°, and 50'. (See fig. 4.) The wing leading-edge two different facilities. The vortex flap configurations slots were sealed for the vortex flap configurations.

tested were a constant-chord full-span flap at three For the purposes of this study, the effects of the flaps deflection angles: 30°, 40°, and 50'. (See fig. 4.)

on airplane mass and inertia characteristics were Static force and moment data were obtained with neglected.

a 0.15-scale model in the Langley 30- by 60-Foot Tunnel at a dynamic pressure of 4.0 psf, which Description of Simulator I

corresponds to a Reynolds number of 1.3 x lo6

The Langley Visual/Motion Simulator (VMS) is based on the wing mean aerodynamic chord. Angle a six-degree-of-freedom motion base simulator. The of attack was varied from -3' to 51° and angle VMS is equipped with a generic fighter cockpit. The of sideslip was varied from -20' to 20'. Dynamic flight instruments available to the pilot include angle forced-oscillation tests were also made about the roll, of-attack and sideslip indicators in addition to the yaw, and pitch axes to determine dynamic stability normal flight instruments. derivatives. A limited set of ground effect data was The control stick and rudder pedal forcefeel char- obtained using a 0.10-scale model in the Langley acteristics were simulated to correspond to the actual 12-Foot Low-Speed Tunnel. Ground effects data were airplane force-feel system. The throttle lever was of obtained for only the basic configuration and the 50' a generic design and was a serial-type control. The vortex flap configuration. Data obtained in the wind- throttle lever on the F-106B is a parallel-type control tunnel tests were evaluated directly by calculating allowing afterburner before military power; however, the force and moment derivatives to determine static this difference was not significant to this study.

and dynamic stability levels of the configurations A closed-circuit color television system provided over the angle-of-attack range tested. Results of the a visual scene of a terrain board that was displayed wind-tunnel tests are reported in reference 9 and will through a virtual image system forward of the front be summarized in this paper.

window. A photograph of the view from the cockpit Comparisons With Handling Qualities during a landing approach is shown in figure 5.

Specifications The VMS is driven by a real-time digital simula- tion system and a Control Data CYBER 175 series Linearized stability and control analyses using the computer system. The dynamics of the simulated full nonlinear six-degree-of-freedom simulation data airplane were calculated by using six-degreeof- base were made to calculate the various response freedom equations of motion at a frame rate of modes of the basic and vortex flap configurations.

32 frames per second. The equations used non- The linear analysis was conducted by operating the linear aerodynamic data as functions of angle of simulation at a given test condition and then lin- attack, angle of sideslip, and Mach number. The earizing the data for small perturbations about the data were obtained from wind-tunnel tests of scale test point. Predictions of response characteristics models and included an angle-of-attack range from were then compared with military handling quali-

t

-3' to 51' and a sideslip range from -20' to 20'.

ties specifications (MIL-F-8785C in ref. 10). In addi- Engine gyroscopic and airplane aeroelastic effects tion, open-loop response simulations using prescribed

i

were not included in the mathematical model.

control inputs were used to investigate control effec- I tiveness and response characteristics. These tests fo- Evaluation Methods cused on the basic and 30' vortex flap configurations because test data (ref. 9) indicated that these con- This investigation focused on the low-speed han- figurations exhibited the highest and lowest levels of dling qualities of the study configurations at the pitch stability, respectively.

critical approach-to-landing flight conditions. As a result, all data shown in this report correspond to Piloted Evaluations those conditions unless otherwise stated. Methods As stated earlier, the pilot-in-theloop evaluations used to study the effects of the vortex flaps on the sta- were focused on the approach-to-landing task. The bility of the airplane included interpretation of static piloting task consisted of an instrument approach ter- and dynamic wind-tunnel test results, comparison of minating with a visual landing. Figure 6 presents results of airplane linear stability analysis with han- a sketch of the task. The simulation was initial- dling qualities specifications, and piloted-simulation ized on a level course offset both vertically and tests.

the basic configuration is 7 percent as compared with laterally from the approach glide slope and local- izer. This forced the pilot to maneuver both lateral- 3 percent for the 30° vortex flap configuration. The reduction in longitudinal stability can be attributed directionally and longitudinally to capture glide slope to two factors: first, an addition of area due to the and localizer while decelerating to approach speed.

The approach was flown in the simulated-instrument vortex flaps ahead of the airplane center of grav- meteorological conditions (no outside view) using ity; and second, the repositioning of the wing vortex "raw" instrument data-glide slope and localizer system.

bars. At an altitude of 250 ft, the airplane en- Static Lateral-Directional Stability tered VFR conditions and the pilot maneuvered The lateral-directional stability characteristics of to a landing. Two research pilots evaluated the the basic and vortex flap configurations are summa- configurations in this study. After each approach rized in figure 10. The data indicate that the basic they were asked to rate the configuration using the configuration remains directionally stable (positive conventional Cooper-Harper scale (ref. 11), shown Cnp) to an angle of attack of 25", and lateral sta- in table 11, and to give any comments on their observations. bility remains positive (negative Cip) to an angle of Several parameters were varied during the inves- attack of 28". The effect of the vortex flap at an- tigation including center-of-gravity location, vortex gles of attack less than 25" is minimal. At higher flap deflection angle, crosswind velocity, and tur- angles of attack, the vortex flaps significantly aug- bulence level. The results of the piloted evalua- ment static lateral-directional stability. The effect of tions were then compared with the analytical data the vortex flaps on control effectiveness is presented obtained and with MIGF-8785C standards (ref. 10) in figures 11 and 12. Figure 11 shows the control when applicable.

moments available using full roll-control inputs for the basic and vortex flap configurations. The data Aerodynamic Characteristics show that the roll-control power is not significantly affected by vortex flaps at angles of attack below As discussed earlier, the vortex flap concept has 30". Figure 12 shows that the vortex flaps do not been shown in previous studies to enhance the perfor- significantly affect the rudder control power in yaw mance of highly swept configurations at maneuver lift at angles of attack below 35". An increased amount coefficients. The improvement in L I D at these con- of rolling moment due to rudder deflection was pro- ditions was also evident in the wind-tunnel data used duced for angles of attack less than 30° compared in this study. (See ref. 9 for additional discussion.)

with the basic configuration-particularly with the The emphasis in the current investigation, however, 30" vortex flap configuration. The data indicate no was on assessing the effect of the vortex flaps on air- plane handling qualities at low speeds. As a result, lateral-directional stability problems in the angle-of- this section will focus only on aerodynamic stability attack range of interest in the approach-tdanding and control characteristics. flight phase.

Dynamic Stability Static Longitudinal Stability Dynamic stability data were obtained using a The longitudinal characteristics of the basic con- conventional forced-oscillation technique (ref. 12).

figuration with several elevon settings are presented in figure 7. The pitching-moment data indicate The results of these tests are summarized for the that the basic configuration is statically stable with primary damping derivatives cmq, Elp, and cnr in a static margin of about 7 percent for the refer- figures 13, 14, and 15, respectively. Although the ence center-of-gravity position and maintains suf- data show some effect of the vortex flaps on these ficient control power to trim beyond an angle of parameters at higher angles of attack, there are no attack of about 35". The major effect of the addi- significant effects at the angles of attack of interest tion of vortex flaps on longitudinal stability was a to this study (a < 22").

reduction of static margin. This effect is indicated Ground Effect in figure 8 which summarizes the effect of vortex flap deflection angle on longitudinal stabilit,y, The data The effect of proximity to the ground was investi- indicate that all vortex flap configurations exhibited gated for the basic and 50" vortex flap configurations.

a marked reduction in static margin compared with The longitudinal data obtained in the study were the basic configuration. Figure 9 presents a compar- used in the simulation. Figure 16 shows the ison of the static margin for the basic and 30" vortex results for the basic configuration. The data indicate flap configurations. The data show that at the nom- a slight increase in lift due to ground effect at an- inal center-of-gravity location, the static margin of gles of attack greater than 8'. The pitching-moment data show a corresponding increment of nose-down calculated. Results for the basic and vortex flap con- pitching moment due to ground effect. This indi- figurations are summarized in figures 19 to 21. The cates that the center of pressure of the added lift data in figure 19 show a decrease in short-period fre- due to ground effect is behind the center of grav- quency due to the addition of vortex flaps. This ity. At angles of attack less than 8 to lo’, the lift is effect is due primarily to the decrease in static sta- reduced in ground effect. The pitching-moment data bility with the addition of the flaps as discussed ear- show an increased nose-up pitching moment due to lier. This characteristic is also primarily responsi- ground effect at low angles of attack, and therefore ble for the significant increase in damping ratio as the “suck-down” effect in lift due to ground effect shown in figure 20. Figure 21 summarizes the ef- would not be expected to create a problem in r e fect of the center-of-gravity location on the phugoid tating the airplane to a lift-off attitude. Figure 17 damping ratio for the basic and vortex flap configu- shows the effect of ground proximity on the 5 0 ’ vor- rations. The results are compared with values from tex flap configuration. Similar trends are seen as dis- reference 10, and all values fall well within the range cussed for the basic configuration; however, the mag- of acceptable handling quality criteria. It is inter- nitude of the ground effects is less for the 50’ flap esting to note the rapid increase in damping ratio at configuration. the aft c.g. range for the vortex flap configurations.

The observed “jump” in the damping ratio occurs at the point where the short-period mode becomes Non-Piloted Simulation Results nonoscillat ory.

Analytical studies using the simulation data base An important consideration in the handling qual- were conducted to evaluate the stability and con- ity evaluation of a configuration is the sensitivity of trol characteristics of the basic and vortex flap the controls. Excessive pitch-control sensitivity, com- configurations at approach-to-landing conditions.

bined with low static stability levels, can make an The classic stability and control parameters were cal- airplane prone to pilot-induced oscillations and over- culated for the flight conditions shown in table I11 control. Figure 22 summarizes the pitch-control sen- and were compared with values listed in reference 10 sitivity for a range of center-of-gravity locations for corresponding to desirable handling qualities.

the test configurations. The data show an increase in pitch sensitivity due to the addition of the vortex Longitudinal Stability flaps and due to aft movement of the center of grav- Longitudinal stability characteristics were calcu- ity. The data are compared with values given in ref- lated using linearized analysis methods described erence 10 corresponding to three levels of handling in reference 10. Flight path stability during the qualities. The data show that the stick-forceper-g approach-telanding condition for the basic and 30’ values for the basic configuration are in the satisfac- vortex flap configurations is presented in figure 18. tory region over the entire center-of-gravity range, The figure shows the flight path angle as a function of whereas the values for vortex flap configurations are in the unsatisfactory region over much of the center- velocity. The data were calculated by setting initial conditions on a 3’ glide slope at a nominal approach of-gravity range. The most sensitive configuration speed of 180 knots. The velocity was then varied over most of the center-of-gravity range is the 30’ through pitch-control inputs without changing the vortex flap configuration which does not meet the power level. The results show that the basic config- sensitivity levels for satisfactory handling qualities uration is slightly unstable at the normal approach at any point in the c.g. range.

speed of 180 knots. This indicates that the basic con- The major factor causing the increased pitch- figuration is operating on the backside of the power control sensitivity of the vortex flap configurations curve at the approach flight condition. Guidelines is the reduction of static stability discussed earlier.

for the amount of flight path instability allowable However, it should be noted that even at comparable during the approach are published in reference 10. levels of static stability, the vortex flap configurations The maximum unstable slope for level 1 flying qual- maintain a higher level of pitch-control sensitivity ities (table 11) is indicated in figure 18. The results than the basic configuration. This characteristic show that characteristics of the basic configuration can be seen, for example, by comparing results at fall within the acceptable range. The 30’ vortex flap center-of-gravity locations of 24 percent E for the 30’ configuration shows flight path stability at 180 knots, vortex flap configuration and 28 percent E for the an indication that it is operating on the front side of basic configuration. In these cases, the higher control the power curve. sensitivity for the vortex flap airplane is primarily Values of the frequency and damping ratio of the due to the increased elevon moment arm resulting short-period and phugoid oscillation modes were also from the forward c.g. movement. The longer moment indicate that the control of sideslip excursions during arm gives increased pitch-control effectiveness and roll maneuvers is satisfactory for both configurations.

results in higher levels of pitch-control sensitivity when compared with the basic configuration at the Pilot Evaluations same level of static stability.

Pilot evaluations were conducted to assess the Lateral-Directional Stability closed-loop stability and handling qualities of the basic and vortex flap configurations. The pilot Lateral-directional stability characteristics of the evaluations concentrated on the approach-to-landing study configurations were investigated in a manner task in simulated-instrument meteorological condi- similar to that employed in the previous longitudinal tions as described previously. This task required study. Linear stability analysis was used to calculate the two pilots (A and B) involved in the study to values of Dutch roll frequency, damping ratio, and maneuver both laterally and longitudinally to c a p the roll and spiral mode time constants.

ture glide slope and localizer. Simulated approaches The Dutch roll characteristics obtained for the were conducted for the basic and 3 0 ' vortex flap basic and vortex flap configurations are presented configurations. A less comprehensive test matrix in table IV. A comparison with the criteria of refer- was evaluated with the 40' and 50' vortex flap ence 10 shown in table V indicates that all the study configurations.

configurations exhibit desirable (level 1) Dutch roll Effect of Center-of-Gravity Location frequency and damping.

The criteria for spiral stability, as shown in refer- As previously shown, the major effect of incor- ence 10, specify that following a disturbance in roll porating the vortex flap is a reduction in pitch angle, the time for the roll angle to double amplitude stability combined with an increase in control sen- shall be greater than 12 sec for level 1 flying quali- sitivity. Studies indicated that at the nominal c.g.

ties. The analysis showed that both the basic and location of 28 percent E, the basic configuration ex- vortex flap configurations were spirally stable and hibited the most favorable handling qualities followed tended to reduce the roll angle after a disturbance, by the 50°, 40°, and 30° vortex flap configurations in although the addition of the vortex flaps decreased order of increasingly degraded controllability. This the level of stability. The calculated time to reach trend is shown in figure 26 in terms of the Cooper- half-amplitude for the basic and 30' vortex flap con- Harper rating scale. The best and worst configura- figurations is presented in figure 23 as a function of tions, the basic and 30° vortex flap configurations, center-of-gravity location.

respectively, were investigated further to determine Roll response characteristics were calculated for the effect of center-of-gravity location on flight char- the basic and vortex flap configurations. Figure 24 acteristics. Figure 27 summarizes the test results for presents time history data of an abrupt input of the two configurations over a center-of-gravity range one-half the available antisymmetric elevon con- from 24 percent C to 32 percent C. Figure 27(a) trol deflection applied at the approach flight con- presents the pilot ratings for the approach task with dition. The resulting roll rate for the basic con- the basic configuration. The ratings indicate that the basic configuration was judged to have fairly good figuration was significantly greater than that for handling qualities at all center-of-gravity locations the 30° vortex flap configuration. These differ- except at 26 percent E . The degraded pilot rating for ences in roll response between the basic and vor- tex flap configurations are due to aerodynamic the airplane with the c.g. at 26 percent E was a re- roll damping and rudder characteristics. The roll sult of the increased difficulty that the pilots experi- rate damping of the vortex flap configurations is enced in maintaining precision control in pitch. Pitch slightly greater than that of the basic configuration oscillations were observed which required consid- (fig. 14), and in addition the rolling moment due erable pilot compensation to make an adequate to rudder deflection is more adverse (fig. 12). approach.

Because of the elevon-rudder interconnect feature in Figure 27(b) presents the results with the 30° the control system, the rudder is deflected with the vortex flap configuration. The pilot ratings show antisymmetric elevon deflection, thus producing a good agreement, between the two pilots and clearly roiiing moment from the rudder deflection that op- indicate a degradation in handling qualities when poses the rolling moment produced by the antisym- compared with the basic configuration. The results metric elevon deflection. Values of the sideslip excur- show that satisfactory characteristics (level 1) were sion parameter A&,/k for the two configurations not achieved at any c.g. location; however, adequate are presented in figure 25 and compared with the performance was obtainable at center-of-gravity lo- handling quality criteria of reference 10. The results cations at and ahead of 28 percent E . As expected, __ based on the wind-tunnel results discussed earlier, are summarized in figure 29. The data show that the lower pilot ratings for the vortex flap configura- artificially reducing the pitch-control sensitivity en- tions were due primarily to degraded pitch character- hanced the handling qualities of the configuration to istics; no significant differences in lateral-directional the extent that an improvement of approximately handling qualities were noted between any of the 1 pilot rating point was obtained on the Cooper- study configurations. With the vortex flaps, the pitch Harper rating scale. The results indicate that re- attitude and vertical speed were difficult to main- stricting the c.g. to 28 percent C combined with in- tain at constant values during the approach. This creasing the force-feel gradient at low speeds provides characteristic was aggravated as the center-of-gravity the vortex flap configuration with flying qualities location was moved aft. It was also found that dur- approaching those of the basic airplane.

ing a maneuver to capture the localizer, a roll in- put would cause a pitch disturbance that increased Wind and Turbulence the pilot work load. This nose-up characteristic was observed in both the basic and vortex flap configu- The effects of wind and turbulence were investi- rations but was much more pronounced for the lat- gated primarily on the basic and 30" vortex flap con- ter. The nose-up disturbance during a roll was found figurations at several center-of-gravity locations. A p to be a result of increased elevon effectiveness in a proaches with crosswind components up to 20 knots trailing-edge-up sense such that a symmetric differ- were conducted. The effect of crosswind below ential deflection for aileron control produced a slight 1 0 knots was very minimal in terms of pilot rating nose-up pitching moment.

for both configurations. At 20 knots, the crosswind component became a significant factor and degraded Control System Modifications the pilot ratings by about 2 points on the Cooper- As previously discussed, the 30" vortex flap con- Harper scale because of difficulties in the flare portion figuration is less statically stable in pitch than the of the landing.

basic configuration. It is interesting to note that The turbulence model employed by the simu- even at center-of-gravity locations at which the same lation was the standard Dryden turbulence model level of stability exists, a discrepancy is still present (refs. 1 3 and 14). Random turbulence in the three in ratings between the configurations. For example, axes was input at light and moderate levels. Turbu- with a center-of-gravity location of 32 percent 3, the lence at these levels was not found to degrade the basic configuration was rated to have satisfactory fly- handling qualities of the configurations significantly ing qualities (PR = 3). With the c.g. at 28 per- during the approach, but some degradation was evi- 3 to achieve the same level of stability, the vor- cent dent during the flare segment for configurations with tex flap configuration was rated to have significantly reduced longitudinal stability.

poorer characteristics (PR = 5 ) . This difference is attributable to the much higher pitch-control sen- sitivity of the vortex flap configuration discussed Conclusions earlier (fig. 22) which caused overcontrol and pilot- induced oscillations during the simulated approaches.

A piloted-simulation investigation was conducted One approach for addressing this problem is to to study the effect of vortex flaps on the low-speed increase the stick force gradient and therefore the handling qualities of a delta-wing airplane. The stick-force-per-g values. The control system of the results can be summarized as follows: F-106B airplane varies the pitch stick force gradient as a function of dynamic pressure during flight. To 1. The reduction in pitch stability from about investigate the effect of decreasing the pitch-control 7 percent to 3 percent stable static margin associated sensitivity for the 30" vortex flap configuration, a with the installation of the vortex flaps degraded the bias in dynamic pressure (designated as "ij-bias") was handling qualities of the airplane in the approach-tw added to the control system to increase the pitch stick landing task by increasing the difficulty in controlling force gradient at low-speed flight conditions. The pitch attitude and rate of descent.

bias was selected such that the resulting control sen- sitivity would be equivalent to that of the basic con- 2. Acceptable handling qualities can be achieved figuration with comparable levels of static margin. for the vortex flap configurations by limiting the aft Figure 28 shows the pitch-control sensitivity for the center-of-gravity location to 28 percent of the mean two configurations across the center-of-gravity range aerodynamic chord. Further improvement can be obtained by modifying the flight control force-feel after the bias was added to the vortex flap configu- ration. The results of the piloted-simulation studies system to reduce pitch-control sensitivity.

5. Lamar, John E.: Subsonic Vortex-Flow Design Study 3. Changes in lateral-directional handling char- for Slender Wings. J . Aircr., vol. 15, no. 9, Sept. 1978, acteristics due to the addition of vortex flaps were pp. 611-617.

minimal.

6. Frink, Neal T.: Concept for Designing Vortez Flap 4. The addition of vortex flaps did not signifi- Geometries. NASA TP-2233, 1983.

cantly affect the characteristics of the airplane in 7. Lan, C. Edward; and Hsing, C. C.: Subsonic Analysis ground effect.

and Design of Vortex Flaps. Vortez Flow Aerodynamics, Volume IZ, James F. Campbell, Russell F. Osborn, and NASA Langley Research Center Jerome T. Foughner, Jr., eds., NASA CP-2417, 1986, Hampton, Virginia 23665-5225 pp. 97-110.

July 29, 1987 8. Teper, Gary L.: Aircraft Stability and Control Data.

Contract NAS2-4478, Systems Technology, Inc., Apr. 1969. (Available as NASA CR-96008.)

References 9. Yip, Long P.: Wind-Tunnel Free-Flight Investigation of a 0.1 5-Scale Model of the F-l06B Airplane With Vortez 1. Rao, Dhanvada M.: Leading Edge Vortez-Flap Ezperi- Flaps. NASA TP-2700, 1987.

ments on a 74 Deg. Delta Wing. NASA CR-159161, 1979.

10. Military Specification-Flying Qualities of Piloted Air- 2. Hallissy, James B.; Frink, Neal T.; and Huffman, planes. MIL-F-8785C, Nov. 5 , 1980. (Supersedes MIL- Jarrett K.: Aerodynamic Testing and Analysis of Vor- F-8785B, Aug. 7, 1969.)

tex Flap Configurations for the 5-Percent Scale F-106B.

The 11. Cooper, George E.; and Harper, Robert P., Jr.: Vortez Flow Aerodynamics, Volume II, James F. Camp- Use of Pilot Rating in the Evaluation of Aircraft Handling bell, Russell F. Osborn, and Jerome T. Foughner, Jr., NASA TN D-5153, 1969.

Qualities.

eds., NASA CP-2417, 1986, pp. 227-248.

12. Chambers, Joseph R.; and Grafton, Sue B.: Investiga- 3. Yip, Long P.: Investigation of Vortex Flaps on the tion of Lateral-Directional Dynamic Stability of a Tilt- F-106B Airplane Configuration in the 30- by 60-Foot Wing V/STOL nansport. NASA T N D-5637, 1970.

Wind Tunnel. Vortez Flow Aerodynamics, Volume II, 13. Taylor, James: Manual on Aircraft Loads. AGARD- James F. Campbell, Russell F. Osborn, and Jerome T .

ograph 83, Pergamon Press Inc., c.1965.

Foughner, Jr., eds., NASA CP-2417, 1986, pp. 201-226.

14. Houbolt, John C.; Steiner, Roy; and Pratt, Kermit G.: 4. Erickson, Gary E.: Application of Free Vortex Sheet Dynamic Response of Airplanes to Atmospheric n r - Theory to Slender Wings With Leading-Edge Vortex bulence Including Flight Data on Input and Response.

Flaps. AIAA-83-1813, July 1983. NASA TR R-199, 1964.

c

Table I . Mass and Dimensional Characteristics of Simulated Airplane

Weight. lb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27400

Moment of inertia:

Ixx. slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15 983

I y y . slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174383

I z z . slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184 002

1x2. slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5971

Wing:

Span. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38.29

Area. ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 695

Mean aerodynamic chord. ft . . . . . . . . . . . . . . . . . . . . . . . 23.76

d N p?

W I-z 0 - d LD UJ h 1 l - ~a a c r U a , L a .- W I - U I 6 V a l v) W L aJC -1 L O w z c

.- .-

0 0 a 0 H .r U r d ZI- 0 w a J v ) ++a a v L e cr m c a l I - W e a a J a o n C U E 10 a L

:E

c ( E O a v a n 0 % d W V L E 0 W W aJ 0 0 a , x - 0 a IC- V l-3 L .- c O - u a J a 0) z w .- a a L O W a L a J a , .- a J > IC mcr -In c, *- aJ n o aaJ a m L z E U a c

a u .-

U a J c E m C L a J 0 .r u a .- 0 u x n t = I C a m U aJ c a l

.-

IC n a J

m I

l a U V e L CI O a J v) I- m a J . !

c ( W .- > cr a l W I- V =9 a +J W .r a

-

I a U o c .r =9

T-

L L

*

- I

Z n t w a

I

I - W V U .- L 0 w a in z W z m n w *- m W n c - ( C f c r V 0 - W L L 3 n > W z vcr I- O a 3 a c ( 0 0 a W n i i a m a t Table 111. Approach-to-Landing Flight Conditions [Landing gear down; speed brakes out; c.g. = 28 percent 2; Weight = 27400 lb]

I Configuration

30' vortex 40° vortex 50' vortex Flight condition Basic flap flap flap

Velocity, knots . . . . 180 180 180

7, deg . . . . . . . . -3 -3 -3 -3

a, deg . . . . . . . . 11.3 10.6 10.8 11.2

S , , deg . . . . . . . . -4.13 -2.42 -3.05 -2.54

Thrust, lb . . . . . . 6410 5792 5925 5518 Table IV. Summary of Dutch Roll Characteristics of Test Configurations 30° vortex 40' vortex 50' vortex Basic flap flap flap =enter-of-gravity location, WDR, WDR, WDR, WDR, percent 2 rad/sec cDR rad/sec <DR rad/sec (DR rad/sec ~ D R 24 2.11 0.215 1.88 0.282 2.03 0.262 2.13 0.245 26 2.07 .205 1.85 .267 1.99 .255 2.11 .234 28 2.03 .193 1.83 .251 1.94 .249 2.09 .224 30.5 2.00 .175 1.80 .239 1.88 .244 2.07 .210 Table V. Minimum Dutch Roll Frequency and Damping as Specified in Reference 10 Minimum values of- Flying quality ~ D R WDR, W ~ ~ , level ~ D R rad/sec rad/sec 1 0.08 0.15 1.0 2 .02 .05 .4 .4 3 0 B a s i c l e a d i n g e d g e

\

With Figure 1. Vortex flap concept.

E=23.76 f t

e

p-38.29 ft -

N o t e : C e n t e r of g r a v i t y s h o w n a t nominal 27.5% C l o c a t i o n 71.07 f t - 4 S p e e d b r a k e Figure 2. Drawing of basic configuration.

rc v) P 0 ' I 1 II m u 0 0 0 0 0 0 0 0 0 0 0

a w t m c u ' cu m t

. % n

E % - .-

'""c 160

Rig Rudder pedal force, Ib Le L e f t +Right Rudder position, deg (b) Directional characteristics Figure 3. Continued.

Right I Lateral stick force,

t o

Ib

I I I I I I I

2 4 6 8 10 12 14 1 4 1 2 10 8 6 4 2

Left stick -+- Right stick

Differential elevon, deg (c) Lateral characteristics.

Figure 3. Concluded.

C o n f i g u r a t i o n

40°

v o r t e x f l a p

50'

v o r t e x f l a p Figure 4. Sketch of vortex flap configurations tested.

Figure 5. View from cockpit during final approach.

3 u

\

I

\ \ \

-

0 1 I

I i

a

I LL

>

c9 I

/ !

I

L E

I

l n -!

-!

P UE

II .o" c i .I U v) .I L a , U

$

c d c % I 4 2 .- t L d + d d .I U a , % -u a d e : X a , i.2

E

> ru U

&

w od E I tL .I Lr

Baseline

S t a t i c

margin,

% E

-51 I I 4 1 I ]Unstable

2 2 2 4 2 6 2 8 30 3 2

C e n t e r o f g r a v i t y , % E

Figure 9. Static margin of basic and 30' vortex flap configurations.

C Y B -.01 -.02 .004 cnB -.004 -.008 Configuration

- . 0 1 2 U o Basic

.006 .004 .002

C

-.002 -.004 -.006 -.008 -5 0 5 10 15 20 25 30 3.5 40 45 50 55

a, deg

Figiire 10. Effect of vortex flap deflection on lateral-directional characteristics. 6 , = 0'.

.oo 5 c c\ Configuration

-.005 -

Basic

ACn - . O l O - _--- 3 0 ' v o r t e x f l a p

-- 40' v o r t e x flap

-

-.O 15 -- - 5 0 ' v o r t e x f l a p

-.020 L .005 -.005 A C l -.010 -.015 -.020 Figiire 1 1 . Effect of vortex flap deflection on aileron control power. 5 , = 7"; 5 , = 0" .004 -.004 -.008 Configuration

-.o 12

Basic 3 0 ' v o r t e x f l a p -.O 16 40' vortex flap -.020 5 0 ' v o r t e x f l a p -.024

.o - O I T 12 I

--

- . 0 0 4 I I I 1 I I I I I I I I I

-5 5 15 25 35 45 55 a, deg Figure 12. Effect of vortex flap deflection on rudder control power. 6 , = 25".

-

Cmqv rad- Configuration 0 30° vortex flap

0 40' vortex flap

A 50' vortex flap 0 10 2 0 30 40 50

a, d e g

Figure 13. Effect of vortex flap deflection on pitch damping.

-

C l p Configuration rad- 0 Basic 0 30° vortex flap

-1.0 -

0 40' vortex flap

A 50' vortex flap

I I I I

-1.5 0 10 20 30

a, d e g

Figure 14. Effect of vortex flap deflection on roll damping.

Configuration 0 Basic 0 30° vortex flap

0 40° vortex flap

A 50° vortex fl I ‘ - 1 Figure 15. Effect of vortex flap deflection on yaw damping.

I I

c

.09 .08 .07 h/b .06

o a l

.05

0 .491

Cm

.04

0 .392

A .261

.03

b .174

-02

.o 1

.8 .7 .6 .5 .4

CL

.3

.2 cD .1

-. 1

-.2 -.3 -5 0 5 10 1 5 ' 2 0 25

a, deg Cm

Figure 16. Ground effects for basic configuration.

.08 .07 .06 .05 h/b

Cm -04 000

0 .491

.03

0 .392

A .261

.02

.o 1

-

.7

-

.6

-

.5

-

.4

-

.3

-

.2

CL

-

. 1

cD

0-

-

-.l

-

- . 2

-.3

- 4

-

-.4

-

-.5 1 1 1 1 1 I I I I I I 1 I

-5 1

5 10 15 2 0 25 .08.07.06 .05.04.03.02 .O 1 1

Figure 17. Ground effects for 50' vortex flap configuration.

Configuration c - 1 0 Basic 0 30' vortex flap Flight path

-2 c

angle, Maximum positive deg slope for l e v e l 1

-4 I I I

1 7 0 1 8 0 1 9 0 Velocity, knots Figure 18. Effect of vortex flap deflection on flight path stability.

Configuration 0 Basic 0 30° vortex flap 2.0 I-

0 40' vortex flap

A 50° vortex flap n w S p ' rad/'sec

-

.4 I I I I I Figure 19. Effect of vortex flap deflection on short-period frequency.

Configuration 0 Basic

s 8 r

0 30° vortex flap

0 40' vortex flap

soo vortex flap

L P .5 .4 .3 2 2 2 4 2 6 2 8 30 3 2 Center of gravity, % E Figure 20. Effect of vortex flap deflection on short-period damping ratio.

Configuration 0

r

0 Basic 0 30° vortex flap

0 40' vortex flap

A 50' vortex flap P

-

.7

-

.6 (P . 5 -

-

.4

-

.3

- 2 -

.l- Level 1

- - - - - - - - - - - - - - - - - - - - - - --

I I I 1 Level 2 I O b Figure 21. Effect of vortex flap deflection on phugoid damping ratio.

Configuration 0 Basic 0 30° vortex flap

0 40' vortex flap

A 50' vortex flap Satisfactory Stick v

----------

force

4 r$r. cc e D t a b le b ut w a unt s im D r o v e me nt

Per 9 , Ib/g Unsatisfactory 24 25 26 27 28 29 3 0 3 1 Center of gravity, % C Figure 22. Effect of vortex flap deflection on pitch-control sensitivity.

Configuration 0 Basic 0 30° vortex flap Time to half-amplit ude, sec 1c C 1 I I I I C 26 28 30 32 2 24 Center of gravity, % E Figure 23. Effect of vortex flap deflection on spiral stability.

Configuration

Basic

---- -

3 0 ' v o r t e x flap

100-

PI

deg/sec

0 -

- 1 o L

-

0 2 4 6 8 1 0 1 2

t, sec

Figure 24. Response to abrupt roll-control input.

Configuration 0 Basic 0 30° vortex flap

a

c

0 -40 - 1 2 0 -200 - 2 8 0 -360

wp, deg

Figure 25. Sideslip excursion parameter.

1 0 Pilot B

a

Level 2 Pilot rating 5

4 5 00

40° Baseline 30Q v o r t e x f l a p v o r t e x f l a p v o r t e x f l a p Figure 26. Effect of vortex flap deflection on handling qualities. c.g. = 28 percent C.

cu

I ‘cr) I,“

m

c

.- Y 2‘ ho CL: N c G z X Q, N Y bb c .d

F Y

I

2‘ Y m B

I l l I I

- .d

-

N D a v t m c u r O c c .e Y

a m

c,c, : 0 0 x

- -

Y

.- .-

.- >

e a

2‘ bo

0 0

d f L 9) Y

cu

e

cu

Q, cr)

I c r )

c Y

0 G

w

cr) G c . l d E I .d Y ho

co

.e 2‘ k ru I bo ce c

s

(0 .e N 2 Q h c d v

I I

rt N - .-,

Configuration

0 Basic

0 30° vortex flap

0 30° vortex flap with ij-bias

Satisfactory

Stick

force

nt

Per 9,

Ib/g

2 - Unsatisfactory

0 1

24 25 26 2 7 28 29 30 31

Center of gravity, % e

Figure 28. Effect of increased stick-force gradient on pitch-control sensitivity.

1 %

G

I J - m

I

I

I a

-cu

#

I

A .- a

I

h P v

I

- < o v)

cu

e

Q

.-

Q)

a

t

I

1 0

w

1 I I I

-cu

r

Y

-

O Q ) 0 0 r c < o m " d g c u ~ O

.-

C 0 0 0 0 c9m

0 0

IJ

I

I

I

I

I

I

I

I

I I

I I

I I

I I

pj

I 1 I I I dc9 (u 7 0 O Q ) 0 0 b < o r n ~~ Report Documentation Page National Aeronautics and Space Adm#n#Slral~on L . Report No. 2. Government Accession No.

3. Recipient’s Catalog No.

NASA TP-2747 I. Title and Subtitle 5. Report Date Piloted-Simulation Study of Effects of Vortex Flaps

I September 1987

on Low-Speed Handling Qualities of a Delta-Wing Airplane 16. Performing Organization Code r. Author(s) 8. Performing Organization Report No.

Jay M. Brandon, Philip W. Brown, and L-16307 Alfred J. Wunschel 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Paper National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, DC 20546-0001 15. Supplementary Notes Jay M. Brandon and Philip W. Brown: NASA Langley Research Center, Hampton, Virginia.

Alfred J . Wunschel: U.S. Air Force Systems Command, assigned to Langley Research Center, Hampton, Virginia.

16. Abstract A piloted-simulation study was conducted to investigate the effects of vortex flaps on low-speed handling qualities of a delta-wing airplane. The simulation math model was developed from wind- tunnel tests of a 0.15-scale model of the F-106B airplane. Pilot evaluations were conducted using a six-degree-of-freedom motion base simulator. The results of the investigation showed that the reduced static longitudinal stability caused by the vortex flaps significantly degraded handling task. Acceptable handling qualities could be achieved by limiting qualities in the approach-telanding the aft center-of-gravity location, consequently reducing the operational envelope of the airplane.

Further improvements were possible by modifying the flight control force-feel system to reduce pitch- control sensitivity.

17. Key Words (Suggested by Authors(s)) 18. Distribution Statement Vortex flap Unclassified-Unlimit ed Handling qualities Piloted simulation F-106B Subiect Cateaorv 08 ~ 19. Security Classif.(of this report) 20. Security Classif.(o his page) -r21. NO. of Pages 122. Price

Unclassified I Unclassified I A03

I 36

NASA-Langley, 1987 NASA FORM 1626 OCT 86 For sale by the National Technical Information Service, Springfield, Virginia 22161-2171