Application of Wind Tunnel Free-Flight Technique for Wake Vortex Encounters

NASA Technical Paper 3672

Application of Wind Tunnel Free-Flight

Technique for Wake Vortex Encounters

Jay M. Brandon, Frank L. Jordan, Jr., and Robert A. Stuever Langley Research Center • Hampton, Virginia Catherine W. Buttrill Unisys Corporation • Hampton, Virginia National Aeronautics and Space Administration Langley Research Center • Hampton, Virginia 23681-2199

November 1997

Available electronically at the following URL address: http://techrep°rts'larc'nasa'g°v/ltrs/ltrs'html Printed copies available from the following: NASA Center for AeroSpace Information National Technical Information Service (NTIS) 800 Elkridge Landing Road 5285 Port Royal Road Linthicum Heights, MD 21090-2934 Springfield, VA 22161-2171 (301) 621-0390 (703) 487-4650 Defining a safe, operationally satisfactory wake Summary encounter for a given aircraft pair under any one condi- tion has proved very difficult. In addition to being able to A wind tunnel study was conducted to determine the adequately model the decay and advection of a wake vor- feasibility of using the free-flight test technique to study tex in the atmosphere, an equally important element is a wake vortex encounters. A generic business-class jet air- valid model that represents the wake vortex encounter plane model was instrumented and flown in the vicinity itself. Substantial analytical research has been conducted of a wake vortex generated by a simple wing. The in vortex modeling (refs. 2 through 4), in vortex-airplane strength of the vortex could be varied by adjusting the interaction, and in resulting forces and motions (refs. 5 generating-wing angle of attack. The variation in the through 7) to try and quantify or predict hazards. Experi- strength of the vortex allowed researchers to study a mental research has been conducted both in flight (refs. 8 range of simulated vortex strengths for a fixed-span ratio through 10) and in wind tunnels (refs. 11 and 12) to pro- of 0.75 and enabled the simulation of various following vide information on the wake vortex flow fields pro- distances and generator airplane weights without the duced by various airplanes and to investigate loads uncertainties in vortex decay and atmospheric effects.

imposed by an aircraft wake vortex on following The study showed that the free-flight test technique was a airplanes.

viable and useful tool in the study of the wake vortex encounters----combining vortex flow fields, airplane The free-flight test reported in this paper extends the dynamics, sensors, and flight control aspects.

research database by experimentally investigating the dynamic response characteristics of a follower aircraft Data obtained during this test included qualitative during wake vortex encounters. This was the first attempt and quantitative results. Steady-state limits of controlla- to conduct such tests in a wind tunnel, and was viewed bility were documented as a function of vortex strength.

primarily as a feasibility test to determine whether the By flying several vortex encounter trajectories at high free-flight test technique was a useful research tool in vortex strengths, a mapping was conducted of roll angle, wake vortex encounter research. Specific objectives for roll rate, lateral velocity, and vortex-induced roll-rate this test were (1) to see if the model could be flown acceleration. The data quantified the effects of the model safely and maneuvered accurately in a wake vortex flow entering vortex flow fields of varying strengths.

field of specified strengths, (2) to develop photogramme- try techniques for measuring the position of the model Introduction relative to the vortex, (3) to estimate rolling moments imposed on the model due to the vortex flow field while flying, and (4) to conduct exploratory qualitative evalua- The National Aeronautics and Space Administration (NASA) is conducting research that will enable safe tions of relative upsets for various encounter trajectories.

Related work involved selecting appropriate flight con- improvements in the capacity of the air transportation system. As part of this research, the Terminal Area Pro- trol system approaches to enable these tests to be con- ductivity (TAP) program has the goal of safely achieving ducted successfully. Flight test data obtained in the past capacity levels during instrument meteorological condi- were very difficult to analyze and apply in any general tions equivalent to those currently achievable under sense due to large uncertainties in the data. For example, visual meteorological conditions. One element of TAP, two significant uncertainties are the strength and position of the vortex which is encountered. It is well-known Reduced Spacing Operations (RSO), focuses on both lat- (ref. 13, for example) that atmospheric effects play a pre- eral and longitudinal separation requirements. A key dominant role in the dissipation characteristics of a wake concern for reducing spacing requirements, especially vortex. Also, it is very difficult in flight experiments to when an airplane is following large aircraft on an repeatably conduct the same encounter flight paths which approach, is the danger of upsets generated by the wake would allow high confidence in the results. In proposing vortex of the preceding aircraft. Consequently, part of the current NASA research effort focuses on developing and this wind tunnel experiment, it was assumed that the vor- tex characteristics in the wind tunnel would remain validating the technologies required for an automated Aircraft Vortex Spacing System (AVOSS, ref. 1), which essentially constant with time (only be affected by turbu- lence levels in the tunnel and temperature-pressure varia- will properly select safe separation distances for different weather conditions based on the aircraft pair and tions during the day). Additionally, the level of the predicted-measured vortex behavior. Although the vortex strength may be easily controlled by increasing or AVOSS will generally attempt to space aircraft to avoid decreasing the lift of the wake vortex generator. In this any wake encounter at all, it must be able to select a sep- way, the effects of vortex strength on the dynamics of an aration distance at which a wake encounter is safe and airplane can be studied. The results of this study can be operationally satisfactory, should one occur. applied to a range of generating airplane pairs (with span

ratiosof 0.75) andseparation distances bycalculating or

t30 time required to reach 30 ° bank angle

measuring thevortexintensity at thefollower-airplane

free-stream wind tunnel velocity, ft/sec location.

Vso stall speed in landing configuration, ft/sec

Themetric defining asafe andoperationally satisfac-

x downstream distance measured from the

tory vortexencounter mayentailverysmallallowable

generating-wing quarter-chord

perturbations in aircraftattitude, as well as relatively

little required corrective controlactivity.However, in

Z vertical distance, ft

light of thewakeencounters reported herein thatwere

angle of attack, deg

beingflownaspartof a feasibilitytestto evaluate the

free-flight technique, it is important to notethatthesub-

O_ v = 0.5(ct s - Ctp), deg

sequent upsets often anddeliberately exceeded what

angle of sideslip, deg

mightberegarded asacceptable. Exceeding whatwere

F vortex strength, ft2/sec

regarded asacceptable upsets wasnecessary todetermine

thefeasibility limits for planned follow-on testsandto

aileron deflection, _as - Sap

establish howanaircraft will respond towake turbulence

2 , deg

andthemagnitude of controls thatmustbeprovided to

correct forit.Similarly, although theresults fromthistest

_e elevator deflection, deg

maybecombined withresults fromother research efforts

5r rudder deflection, deg

to identifywakeencounter metrics, it should be noted

pitch angle

alsothattheintentofthisexperiment wasnottoestablish

anacceptable levelforanupset metric.

radial location of follower model relative to 0v vortex core location, deg

Symbols

a density ratio of air at full-scale flight condi- b wingspan, ft tions to model flight conditions at sea level roll mode time constant Ct, lift coefficient lift-curve slope per deg roll angle, deg CL= yaw angle, deg C t rolling-moment coefficient Subscripts: OC t Cl_a = _a' per deg f follower airplane fs full scale Clsr g vortex generator wing = b--_r, per deg p port side c mean aerodynamic chord, ft ms model scale normalized acceleration g s starboard side

lx

roll axis inertia, slug-ft 2 ss steady state lr pitch axis inertia, slug-ft 2 v vortex induced

lz

yaw axis inertia, slug-ft 2 Abbreviations: it horizontal tail incidence angle, deg ARI aileron-rudder interconnect lateral acceleration, g units (positive right) ny ATP advanced turboprop n z normal acceleration, g units (positive up) AVOSS Aircraft Vortex Spacing System P roll rate, deg/sec FAR Federal Aviation Regulations pitch rate, deg/sec q FCS flight control system dynamic pressure, psf FS full-scale value R distance of follower model from vortex core, in. hsw hardware switch MS yaw rate, deg/sec model scale value reference wing area, ft 2 N model scale

NLF natural laminar flow characteristics of the model are shown in table I. The

intent of the investigation was to explore vortex encoun-

RSO

Reduced Spacing Operations ters with the follower model in the high-lift-landing- TAP Terminal Area Productivity approach condition; therefore, the configuration had the trailing-edge flaps deflected 35 °. Model control-surface Model Description and Test Techniques deflection limits were _5 a = +20 ° to -20°; _e = +15° tO -25°; and _r : +20° to -20 °. Horizontal tail incidence The overall test technique used for the free-flight angle it could be varied from +2 ° to -10 ° to provide an experiment is illustrated in figure 1. The primary compo- extended range of pitch trim; however, it was fixed at 0 ° nents are a wing in the forward section of the wind tunnel for the current test. No landing-gear geometry was to generate a wake vortex flow field, a model flying included on the model.

unconstrained behind the wing in and around the wake vortex flow field, and instrumentation required to fly the The model also incorporated full-span Krtiger flaps.

model and to obtain data for analysis. Further details of Previous free-flight test results of the model configured the test setup and conduct will be given herein. The coor- as an advanced turboprop (ATP) (ref. 14) indicated that dinate system used in the tests was based on the location this configuration exhibits an abrupt wing drop and of the starboard vortex of the generating wing. Model autorotative departure against full corrective control at position data presented will be referenced to a radial dis- the stall angle of attack. Static wind tunnel test results tance from the measured vortex core location and an showed that the wing drop was due to an abrupt asym- angle. The angle 0 v is defined as increasing clockwise metric wing stall that produced a pronounced rolling from the horizontal three o'clock position relative to the moment. Additional free-flight tests of the configuration, vortex.

modified to include wing leading-edge devices such as Krtiger flaps, showed that the modified configurations Vortex Generator Wing had acceptable overall flying qualities and no significant stability and control problems, even at post-stall angles The model used to generate the wake vortex was a of attack. Sketches and other details of the full-span rectangular planform aspect ratio 7.0 wing with a span of KriJger flaps that were used on the model can be found in 12 ft, a chord of 1.71 ft, and a NLF(1)-0215F general avi- reference 15.

ation airfoil section. The wing was constructed of fiber- glass and epoxy with an aluminum spar. Further details The model was powered with two thrust tubes regarding the airfoil section can be found in reference 14.

installed on each side of the aft part of the fuselage. The Overall aerodynamic forces and moments acting on the thrust tubes only provided thrust to fly the model, and no model were measured on an internal six-component bal- attempt was made to simulate thrust characteristics of ance, in part to enable estimation of the subsequent shed any specific airplane configuration.

vortex strength from lift but also to monitor loads on the test rig. Smoke generator tubes using heated propylene Scaling Discussion glycol vapor were installed along the wing trailing edge to produce smoke at the wingtips to seed the vortex so A straight, unflapped wing was selected to be that it would be visible for the flight encounters. The fig- the wake vortex generator because of the properties ure 2 photograph shows the wing installed in the Langley of rapid vortex rollup and essentially constant vortex 30- by 60-Foot Tunnel (with the smoke generators oper- strength for long downstream distances (ref. 16).

ating). During the free-flight vortex encounters, all avail- Because of these characteristics, vortex strength at the able smoke was concentrated in the starboard vortex to location of the follower model could be estimated for an improve visual definition of the vortex core location. The elliptically loaded wing by using the approximation: figure 2 photograph also shows illumination of the vorti- 2CLU,,_Sg ces by a laser light-sheet technique, which will be F - The vortex core size may not have been l_b g described subsequently in the Free-Flight Tests section.

representative of full-scale flight conditions in these tests. This core size is currently believed to be of minor Follower Airplane importance due to the large ratio of wingspan to vortex The geometric characteristics of the follower model core diameter both in flight and in these wind tunnel airplane are depicted in figure 3. The model was tests.

constructed of fiberglass and epoxy. The mass and To scale results from the current test and relate them geometric properties were scaled to simulate a represen- tative business-commuter aircraft for the purpose of to a full-size airplane, the flow angularity distribution determining flight characteristics from free-flight tests in caused by the vortex on the current follower model and the Langley 30- by 60-Foot Tunnel. Geometric and mass the full-scale airplane must be equivalent. Additionally,

the followermodel musthaveappropriate dynamically

resulting model motions were measured, and pilot

scaled mass andinertiavalues andsimilaraerodynamic

comments were recorded for each flight condition. All characteristics compared tothefull-scale airplane.

vortex encounters were conducted at a tunnel speed cor- responding to 1.3Vso with the airplane flaps configured

Thescaling considerations dictate thattheresults are

for landing.

dependent on wingspan ratio between the generating

During the free-flight tests, the model was attached

wing andthe followeraircraft.To applythe results

to an umbilical chord which supplied pneumatic and

directly,thespan ratiomust beequivalent in flighttothe

electric power and control signals to the model. The

windtunnel testvalues, andthemodel mass characteris-

chord also contained a l/8-in, steel safety cable that was

tics mustbe appropriately scaled. Then,the vortex

controlled by a safety cable operator using a high-speed

strength estimated at thepointof theencounter (atthe

pneumatic winch. The safety cable operator's function

followerairplane location) canberelated to thepresent

was to help launch the model at the start of a test, to

results byusing generating-wing lift coefficients. Forthe

retrieve the model at the end of a test, to keep tension off

current test, thewingspan ratiowas bflbg = 0.75. For dis-

the model from the umbilical cable during the test, and to cussion later in this report, the follower model will attempt to protect the model in an out-of-control situation be considered to be at 0.175 scale, which is representa- by pulling the model out of the airstream.

tive of an experimental test bed aircraft described in reference 17. These factors result in conditions represen- In addition to the safety cable operator, the model tative of a business jet airplane following a commuter flight crew consisted of a pitch pilot, a thrust pilot, and a airline airplane. Table II shows a comparison of roll-yaw pilot. These piloting functions were located in generating-wing lift coefficient, vortex strength, and the positions shown in figure 1 to afford the best view for full-scale vortex strength. The lift characteristics of the controlling the pertinent axes. The separation of the generating wing are shown in figure 4. Due to dynamic piloting duties is very advantageous for several reasons.

scaling relationships, the small-scale models tested with By separating pilots by axes, effective evaluations can be the free-flight test technique develop considerably faster obtained more easily because the pilot is only controlling responses than full-scale airplanes. Some scaling factors the axes he is trying to evaluate. Model control is also are shown in table III.

enhanced by providing the optimal visual perspective for control of each axis. Due to dynamic scaling, the model Free-Flight Tests motions are substantially faster than those of the full- scale airplane, so separation of piloting tasks is beneficial The wind tunnel free-flight tests were conducted in for that reason as well.

the Langley 30- by 60-Foot Tunnel with the technique illustrated in figure 1. With this technique, the remotely The primary component in the free-flight control controlled, dynamically scaled model was flown in the system is a digital minicomputer that was programmed open test section of the Langley 30- by 60-Foot Tunnel.

with the flight control laws (presented in the appendix).

For most of the tests, a vortex generating wing was The computer processed sensor information from the mounted in the forward section of the wind tunnel. As model and command inputs from the pilots to generate previously discussed, the generating-wing angle of attack command signals to drive the high-speed pneumatic was varied to enable selection of an approximate vortex actuators onboard the model. The data sensors on the strength. The vortex core was marked with propylene model included a three-axis rate gyro to measure angular glycol smoke to enable the free-flight pilots to position rates, an accelerometer package to measure normal-, the flying model in desired locations relative to the axial-, and side-force accelerations, a boom-mounted tx/13 wingtip vortex created by the upstream wing. Figure 5 is vane sensor on each wingtip for angle of attack and side- a photograph of the model flying in the vortex during the slip, potentiometers to measure control-surface positions, test. The wind tunnel free-flight tests were used first to and a transducer to measure pressure at the thrust tube for evaluate the flying characteristics of the airplane with the thrust estimations. These sensor data, along with pilot various control laws and then to evaluate the dynamic control inputs, were recorded in the computer for response and controllability of the model near vortices of postflight analysis. Angular rates, linear accelerations, various strengths. Hence, prior to mounting the generat- and tx/_ vane sensor data were filtered with a first order ing wing in the tunnel, the free-flight model was flown to lag filter with a time constant of 0.05 sec before entering evaluate the control law implementation and to adjust the the FCS. Additionally, ot and 13 from the wingtip gains to ensure that the model was well-behaved in free mounted vanes were corrected for angular rates. Post- air. Next, for the vortex encounter tests, the model was flight data reduction included incorporating upwash first flown near the vortex, and then penetrations of the corrections to the tz data based on static wind tunnel test vortex were made from various trajectory paths. The results obtained previously, correction of accelerometer data tothemodel center-of-gravity location, andcalcula- control system is presented in the appendix. This test tionof angular accelerations by differentiation of mea- represented the first time the wind tunnel free-flight test

sured angular ratedata. technique was used to study wake effects on a flying

model. As such, several challenges were overcome to

Additional quantitative flightdatawere recorded by

provide pertinent results. These challenges included fly-

usingphotogrammetry techniques applied to free-flight

ing the model in very precise locations in the wind tun-

testingfor the first time (ref.18).The modelandthe

nel, relative to the vortex, and measuring the model

windtunnel testsection weremarked withphoto reflec-

position and attitude angles in addition to the vortex core

tive spots at known locations. A camera above theexit

location for postflight analysis. Additionally, tests were

cone in thetunnel wasused totrack themodel. Postflight

conducted without the vortex generating wing installed

analysis of thetracker camera datawasusedto obtain

to determine model flying characteristics and to establish

model Eulerangles andposition in thewindtunnel. The

minimum controllable airspeed (Vso), which determined

location ofthewingtip vortex produced bythegenerating

the tunnel speed for the remainder of the testing.

wingalsowaslocated photogrammetrically. Thesmoke

wasmarked in three locations by a laser lightsheet. The

Free-Air Flight Characteristics

forwardpositionwasilluminated by an Argonlaser

located just outside thebalcony usedby thepitchand

The model was easily flown in free air (no vortex

thrustpilots.Themiddleandaft positions wereillumi-

generating wing upstream) in each of the longitudinal

nated bydiode lasers mounted onthewindtunnel ground

flight command modes. The or-command mode was used

board. Theintersection of thesmoke andthelightsheets

throughout most of the free-air flights because it is a tra-

wasthenused todetermine thevortex corelocation dur-

ditional flight control system and pilots are very familiar

ingthetest.Figure 6 is a photograph of thetestsetup

with it. The airplane could be flown with pitch rate

showing thelasers. Thelocations of theplane ofthelaser

damping removed; however, the resulting airplane

lightsheets relative tothequarter-chord of thegenerator

motions were much more lively. The pitch rate command

wingareshown in tableIV. Additional qualitative data

mode was very easy to fly; however, some pilot learning

wererecorded, includingpilot comments and video

was involved for pilots to become comfortable with the

recordings of themodel flights.Table V liststhetrans-

integrated control path, which resulted in the pitch atti-

ducer accuracies anddata system resolutions which were

tude remaining wherever it was when the stick was available forthekeyrecorded parameters during thetest.

released. The g-command system also resulted in an eas-

A discussion ofthephotogrammetry data factors isavail-

ily controllable model.

able in reference 18.

Several options were programmed into the lateral- directional control laws. Similar to the longitudinal case, Free-Flight Test Results some experimentation was conducted to arrive at an acceptable flying airplane for the wake vortex encounter During the wind tunnel free-flight tests, the model was evaluated by using various flight control system fea- task. The unaugmented airplane was found to exhibit tures. Although the flight control system was not consid- unfavorably low damping in roll and therefore required ered a primary area of interest for this test, it was artificial roll damping for acceptable handling qualities.

necessary to provide the pilots with a model possessing Additionally, the aileron-to-rudder interconnect (ARI) feature was used to alleviate adverse yaw and to enable good flying qualities to conduct the vortex encounter the model to fly in the lateral axis with only one control- flights. Three longitudinal flight control system architec- ler (roll). Yaw rate feedback to the rudder was used to tures were developed for the test: or-command, pitch rate reduce Dutch-roll oscillations during flight. Side acceler- command, and g-command. The m-command system was ation feedback to the rudder was flown with and without a direct link between the pitch stick and elevator posi- the ARI engaged and was found to increase the ease of tions and also included pitch rate damping. The pitch rate command system used the pitch stick position to com- maintaining a lateral position in the tunnel; however, it mand a pitch rate, and with no pitch stick input, the degraded the maneuver capability needed to reposition model control laws would attempt to maintain the current the model. Differential angle-of-attack feedback to the pitch attitude. The g-command system used the pitch ailerons from the two wingtip booms was also evaluated, stick position to command normal acceleration (load fac- but as expected, because the o_ signals were rate cor- rected, they did not influence the flight characteristics of tor), and with no pitch stick input, would seek a I g flight the model in free air. After the flying qualities investiga- condition. Additionally, because of the higher rates developed in dynamic model testing, roll and yaw rate tion was complete, the baseline configuration for the lateral-directional control system consisted of an ARI damping augmentation was added to aid in flying the and artificial roll and yaw rate damping.

model in the wind tunnel. More detail about the flight Flight-determined static stability. The lift coeffi- (ref. 19). The g-command and q-command control sys- cient and pitching-moment characteristics could be tems introduced higher order dynamics; however, each deduced in flight by stabilizing at lg flight conditions at system resulted in a model with very desirable handling various tunnel speeds. As the tunnel airflow speed was characteristics for the free-flight task of limited maneu- decreased, the angle of attack of the model increased for vering in free air conditions.

level flight. A comparison of lift coefficient measured in The lateral dynamic response was similarly obtained flight with that measured during a static wind tunnel test both with dampers-on and dampers-off conditions. All of the same model is shown in figure 7. It is important to lateral-directional maneuvers were conducted with the note that the data from reference 15 include pylon- ARI engaged. The roll mode time constant and maxi- mounted engines with propellers generating zero thrust mum roll rates achieved during free-flight handling with an untrimmed model. These data show reasonable qualities evaluations for the bare airframe (dampers off) agreement with data measured on the model of and the augmented configuration (dampers on) are reference 15. Indications of pitching-moment character- shown in table VI. These values also are shown scaled to istics are shown in figure 8 as average elevator angle at representative full-scale values and are compared to the trim angles of attack. Comparison with data calcu- reference 19 requirements for a class II land-based air- lated from reference 15 shows that the model is some- plane during approach. As can be seen, the roll rate what less stable in the current configuration without the damper limited roll rate capability because it did not dif- pylons and propellers. The roll control effectiveness was ferentiate between a commanded or an uncommanded not directly measured in flight, but the static wind tunnel roll rate.

results from reference 15 are shown in figure 9.

Vortex Flow Field Encounters Identification ofl.3Vs,, The airflow velocity in the wind tunnel was progressively slowed to evaluate the After establishing flight control system gains to slow-speed characteristics of the model and to determine ensure good flying characteristics for the model, the vor- the minimum controllable flight speed in the power- tex generating wing was installed in the forward portion approach configuration. The model departed controlled of the wind tunnel test section. Initial flights were made flight due to insufficient pitch and roll control at a with the wing at zero lift angle of attack, and it was noted dynamic pressure of 5.4 psf, which resulted in a model that there was significantly more turbulence experienced angle of attack of approximately 10 °. Based on these by the model due to the wing and its support (fig. 6). An results and Federal Aviation Regulations (FAR), which interesting phenomenon occurred when the generating- allow an approach speed of at least 1.3Vso, the remainder wing angle of attack was slightly increased to generate a of the testing was conducted at a tunnel dynamic pressure small amount of lift. The flow in the tunnel appeared of 9.0 psf, which is equivalent to a full-scale approach much smoother when there was a slight increase of speed of about 120 knots.

generating-wing angle of attack when flying the model.

The wake generated by the wing at low lift coefficients is Dynamic response characteristics. Flight control believed to have smoothed out the turbulence generated system (FCS) gains were adjusted in real time during by the support structure by the time it reached the area in initial flights in the wind tunnel to arrive at acceptable which the model was flying.

flying qualities for the tests. After setting the FCS gains, dynamic response characteristics were converted to As the wing angle of attack was further increased, full-scale values and compared with airplane handling generating wake vortices, the flow field characteristics qualities criteria to ensure that the model was still repre- were very evident in the flight characteristics of the sentative of actual full-scale airplanes. Model dynamic model. Figure 10 shows a schematic of the flow field response characteristics for the various flight control sys- characteristics around the wake vortex. The starboard tem modes were obtained by performing doublet control vortex (right vortex looking upstream) was the vortex inputs and observing the resultant model motions. The used for vortex encounters. The figure shows an area of typical free-flight control system has traditionally been upwash to the right of the vortex and downwash to the an a-command system with proportional feedbacks for left of the vortex (between the starboard and left wing- stability. The or-command system, which provided the tip vortices). This upwash-downwash flow field made it pilot with what he considered to be good characteristics, very difficult to accurately position the model vertically, gave a short-period mode with a frequency of approxi- while maneuvering laterally. Each longitudinal com- mately 1.56 Hz (0.65 Hz full scale) and a damping ratio mand mode was evaluated to determine the best of approximately 0.6. These dynamic characteristics, configuration for conducting the vortex encounters. The along with a n/ct = 5.9 g/rad, meet level I requirements g-command mode minimized the vertical excursions of for a full-scale airplane based on military specifications the model while it was flying through the vortex flow

field.Thiscommand mode wasused in alltransient vor-

lized in the vortex flow field up to a CL° of approxi- mately 0.7. At that point, roll control on tlqe model was texencounters presented herein.

saturated, and the model could not maintain position in

Thevortexgenerated by theupstream wingexhib-

the vortex.

itedits ownsetof dynamics. Thevortexcorelocation

wandered approximately +5 in. in the vertical and lateral

As the vortex in which the model was flying directions, as shown in figure 11. These data represent increased in strength, the flow field sensed by the the vortex core location movement over a time interval of wingtip-mounted ot/[3 vanes showed the increase in rota- approximately 0.5 sec, with the wind tunnel operating at tional flow. Figure 13 shows the difference between the a dynamic pressure of approximately 5.3 psf. The data angle-of-attack measurement at the left and fight were obtained with the free-flight model removed from wingtips. These data were obtained with the model posi- the test section. The vortex core position was measured tioned directly in the vortex core. The data show that as by using photogrammatic techniques described in refer- the vortex strength is increased, the difference in the ence 18 to locate spatially the centroid of the smoke angle of attack of the right and left wingtip increases.

marking the vortex core that was illuminated by laser 3.

The predicted rolling-moment effects on the model due Laser 3 was located beneath the vortex track a distance of to the vortex flow field are shown in figure 14. These x/bg = 2.6. Most free-flight testing was conducted at a data were calculated by using the differential wing angle wind tunnel dynamic pressure of 9.0 psf; however, the of attack and by assuming a linear upwash-downwash meander of the vortex was qualitatively the same at distribution across the span of the model with the follow- either wind tunnel dynamic pressure.

ing equation: A simple calculation was made to estimate the vortex-induced rolling moment on the follower model during vortex encounters. The assumptions required for this calculation were that the control power effectiveness where CLa = 0.107 was obtained from free-flight data was invariant during the encounter (given by the data records. Though the assumption of linear upwash- in fig. 9). Effects of roll rate damping and dihedral downwash distribution across the span of the model does were neglected. Specifically, the vortex-induced rolling not physically represent typical vortex flow-field charac- moment was estimated by teristics, the prediction based on differential _ shows reasonable agreement with free-flight data (fig. 12) as to "_fl X the strength of the vortex in which the model can be CI_ - qSb - Cl_ 8a- Cl_r_r flown in steady-state controlled flight.

Although not a part of the current study, a compari- Transient vortex encounters. Most research flight son between statically measured and dynamically maneuvers were conducted to obtain data as the model derived rolling moments may be useful in addressing the was maneuvered near and through the core of the star- importance of dynamic effects in predicting vortex encounter hazards. board vortex. As previously discussed, several encounter profiles were flown with the generating wing set at vari- Steady-state encounters. A systematic set of vortex ous angles of attack to provide a wake vortex strength encounters were flown to establish steady-state condi- variation. The transient vortex encounters began with a tions in the vortex. These test points were conducted by generating-wing angle of attack of 8" for a CLg of establishing the generating-wing angle of attack and the approximately 0.95 and ended at the maximum wing resulting flow field, and then the model was flown into angle of attack of 13.4 ° that was available in the tunnel the smoke marking the vortex core of the tip vortex. The for a CLg of approximately 1.25. As previously men- model was stabilized with the fuselage in the center of tioned, testing was conducted with vortex strengths far beyond what would be expected to yield acceptable flight the vortex core (marked by smoke) for several seconds encounter boundaries. This testing was done primarily for data recording. The test runs started with the generat- ing wing at zero lift, and then the angle of attack of the for two reasons: (1) to validate and explore fully the use of the free-flight test technique in these applications, and wing was increased in 1° increments until the free-flight model could no longer maintain position in the vortex (2) to obtain a set of data with large amplitude effects for flow field. To aid in the application of these data in a possible use later in the validation of modeling and general sense, all generating-wing vortex strength data prediction techniques. Additionally, the large vortex will be represented by using the wing lift coefficient as strength data ensured that the observed effects were sig- shown previously in table II. Figure 12 shows the result nificantly greater in magnitude than would occur due to normal wind tunnel turbulence.

of the steady-state encounters. The model could be stabi- Time-history data for vortex encounters for a range Pilot comments at a CLg -- 0.95 indicated that the of vortex strengths and for the four general encounter tra- vortex produced rolling moment when the model was jectories are discussed in this section, and overall trends near the vortex core, exceeded the control capability of from the data will be discussed subsequently. The four the model, and usually resulted in uncommanded left roll encounter trajectories were (1) to descend directly rate and translation down and to the left. Recovering con- through the starboard vortex core, (2) to climb directly trol of the model, once it was away from the core loca- through the vortex core, (3) to translate from the right tion, was accomplished easily, and the ability to regain side of the vortex to between the generating-wing vortex positive control of the model before exceeding the wind pair, and (4) to translate from between the vortex pair tunnel test section area was never seriously in doubt.

through the vortex to the right side outside the vortex.

Figures 19 through 21 show data for CL_ = 1.07.

Model position data will be presented relative to the vor- Figure 19 shows a horizontal approach from right to tex position which was measured during the run at laser left. The trajectory shows, by a rise in the flight path, the location 2 at the initiation of a data run. Note that the effect of the upwash as the model approaches the vortex.

actual vortex core location at the follower model likely As the model crosses through the vortex, it develops a would differ due to downstream distance and the interfer- left-wing-down-roll attitude and descends out to the left ence effects of the follower model on the vortex. The of the generating-wing starboard vortex The left wingtip vortex core (marked by smoke) tended to go around the probe dramatically shows the crossing of the vortex by model rather than to impact directly; therefore, it was the large upwash just prior to the wingtip entering the very difficult to position the model directly in the vortex vortex core and by the immediate change to large, center. The sign convention for the position (radius and negative-sensed angles of attack after the wingtip crossed angular position) is shown in figure 15. Most of the free- through the vortex core. The distance to the vortex core flight data were obtained at a downstream distance from at which a begins to increase (t = 1.3 sec) is approxi- the generating-wing quarter-chord location of approxi- mately 80 in. from the model center of gravity or 26 in.

mately 24 ft (2 spans of generating wing). Figure 16 from the wingtip. At this location, the vortex flow field shows a histogram of the distance between the generating produces a small, positive rolling moment (Cry > 0).

wing and the follower model, including all data points The model center of gravity distance to the vortex is reported herein.

29 in. before vortex-induced rolling moment to the left (Clv < 0) is produced. As seen for the CL 8 = 0.95 case, Figures 17 and 18 show data for a vortex strength the model controls are saturated against the vortex- corresponding to CL_ = 0.95. Figure 17 shows a horizon- induced left rolling moment as the model flies through tal approach from right to left. As the model crosses the vortex core. The large roll angle, to which the model through the starboard vortex core (t = 5.5 sec), the vortex was disturbed due to the vortex, elicited pilot comments flow field causes the model to descend and roll to the that this condition resulted in less certainty of easily left. The time-history data show full-right lateral controls regaining control after the vortex encounter; however, to oppose the vortex-induced moments while the model the control was regained without exceeding the wind tun- is near the vortex core. As the left wingtip approaches the nel test envelope limitations.

vortex, the flow vanes on the wingtip boom show a large upwash and also substantial sidewash. As the wingtip Vortex penetration attempts were also flown by passes beyond the vortex core location, an abrupt change using descending trajectories (fig. 20) and climbing tra- from upwash to downwash occurs. After crossing jectories (fig. 21). During the descending trajectory, the through the vortex, pitch attitude increased to enable pilot had difficulty in positioning the model directly over level flight in the downwash flow field that existed the vortex core due to the effects of the rotational vortex between the two vortices created by the generating wing.

flow field. The trajectory shows that the model passed to the left of the vortex position rather than through the core Figure 18 shows a vortex penetration attempt begin- as intended. Roll control and roll rate activity show the ning between the vortex pair and translating to the right. difficulty experienced by the pilot in attempting to posi- Despite nearly full-right controls, less than 10 ° of right tion the model. When the model approached the vortex, bank could be generated, and thus the translation rate full controls were used again to oppose the vortex- was very small. Even with full-right controls while near induced rolling motions while the model was rolled and the starboard vortex, the model was rolled to the left, was descended out the vortex to the left. The climb through unable to pass through the vortex, and instead translated the vortex (fig. 21) shows that the model again failed to back to a lateral location between the vortex pair. This pass directly through the reference vortex core location, maneuver was repeated at a faster lateral translation rate although it was very close. Again, full opposing controls and resulted in a successful transition through the vortex were required, and the model was pushed away towards to the free air on the right side of the tunnel.

the left of the vortex as it climbed. The wingtip probe did not indicatethe presence of the vortex prior to the pushed out rather violently to the left and down. Full- encounter for the vertical trajectories. Pilot comments right lateral control was required, beginning when the model was within 29 in. of the reference vortex position.

indicated that the vertical trajectories were much less The wingtip probe shows that the wing experiences dramatic than the horizontal trajectories for this effects of the left generating-wing vortex flow field at the generating-wing angle of attack because of the larger beginning of the run. Also note that the model pitch atti- bank angles generated from the horizontal entry, and because the controls were set to substantial left roll com- tude is much higher (9 ° versus 3 °) for the case in which the model is initially between the generating-wing vortex mands in order to penetrate the vortex, which then had to be rapidly reversed when the model had reached nearly pair. The increased pitch attitude is an indication of the into the vortex core. For the vertical entries, large left- downwash field between the vortices, and pilot com- roll control was not needed, and the resultant roll angle ments noted difficulty in climbing while in this location.

perturbations were less.

Vertical trajectories through the vortex were flown Figures 22 through 25 show data for a vortex with both descending and climbing approaches (figs. 24 and 25). For the descending approach, it was very diffi- strength corresponding to CLg = 1.18. Figure 22 shows a horizontal approach from right to left. As the model cult to position the model to result in a good penetration approached the vortex, the upwash flow field generated a of the center of the vortex, as had been seen at the previ- positive (roll-away) moment as indicated by positive val- ous lower vortex strengths. Additionally, recovery was complicated by the lack of climb performance capability ues of CI, and required the lateral control deflections.

The maximum value of positive rolling moment occurred when the model was positioned between the generating- when the model center of gravity was approximately wing vortex pair. Figure 24 shows a high-to-low vortex 50 in. from the vortex. The sign of the vortex-induced penetration. The trajectory shows that the model was ini- moment changed at approximately 29 in. and reached a tially pushed slightly to the left; then, right controls were maximum value at the minimum distance, that is, model applied to attempt an intersection with the vortex core center of gravity in the vortex core. Controls were satu- and fuselage. Typically, in these flights, the model still slightly missed flying directly through the vortex core.

rated for a substantial time during the penetration to Other than the positioning problems, the encounter was oppose the left roll generated by the vortex. The wingtip similar to what had been experienced at the previous vor- probe showed the effect of the vortex with an increase in tex strengths. Figure 25 shows a low-to-high vortex pen- upwash that was measured beginning at a distance of 118 in. (t = 0.9 sec) and then a further rapid increase etration. The trajectory of the model shows a large beginning at about 80 in. (t = 1.8 sec). The flow changed perturbation to the left after the encounter. Roll controls were saturated for a long period. This penetration geome- to downwash again as the wingtip passed the vortex posi- try was repeated several times, and some of the penetra- tion. Sideslip for this encounter showed a rapid increase beginning at about 62 in. (t = 2.1 sec) from the reference tions resulted in the model being recovered by the safety vortex position and changed sign to negative sideslip as cable operator after going out of control and exceeding the wingtip crossed the vortex. This sideslip response the wind tunnel test section envelope. Pilot comments indicated that control of the airplane through the vortex differs from the results shown at CL 8 = 1.07 where the was much more difficult, and confidence was low on the sideslip initially showed negative values as the vortex ability to retain control of the model. Another interesting was approached from the right. Slight differences in the note is that control seemed more in question at this vor- vertical positioning of the model, relative to the vortex, tex strength for vertical penetrations, whereas at lower may be responsible for this measured difference. Addi- vortex strengths, it was felt that the lateral penetrations tionally, as previously noted, the vortex tended to move were more hazardous.

around the model when the model approached the vortex.

This motion is not reflected in the plotted data because Figures 26 through 28 show data for vortex encoun- positions were recorded relative to the reference vortex ters with CLg = 1.25. Several attempts were made to location.

cross the vortex from right to left. Figure 26 shows one of these penetration attempts. As the trajectory plot Vortex encounters also were flown from left to right, shows, the model did not pass through the vortex. The and pilot comments indicated that the model was more upwash field on the right side of the vortex induced more difficult to position than for right-to-left translations.

Unless a reasonable translation rate was established, it right roll than was available through the controls. The data showed full left lateral controls while the model was was impossible to overcome the left rolling moment and successfully pass through the vortex. Figure 23 shows an moving away from the vortex to the right. The wingtip example of an unsuccessful penetration attempt from the probe showed that it apparently crossed the vortex core left. The trajectory shows the model translating from left briefly during the encounter attempt with the large to right, and when the model just reaches the vortex, it is increase in upwash followed by downwash when the Note that the percentage of roll control power available

wingtip probe wentthrough thevortex core. Sideslip data

in figure 29 is estimated based on aileron effectiveness

showed a largenegative sidewash approaching thecore

measured in a static wind tunnel test outside the influ-

and positivesidewash oncethe wingtip probewas

ence of the wake vortex system.

throughthe core. Pilot comments andvideorecords

showed thatit wasveryhardtointersect thevortex core

To identify systematically the effects of flight duetomodel interaction withthechanging upwash field.

through a wake vortex as the vortex strength was

Additionally, largetranslational rateswererequired to

increased, data were obtained from all vertical trajectory

successfully penetrate throughthe vortexbecause of

vortex encounters. The horizontal encounter data were

insufficient lateralcontrolpowerto oppose thevortex-

not included due to the large amounts of scatter attribut-

generated moments. Evenwiththefaster rates, when the

able to differences in pilot technique between runs. One

model wasflownjustbeloworjust above thevortex, as

proposed measure of identifying wake vortex hazard is marked bythesmoke, it could translate across thevortex; by quantifying maximum allowable bank angle excur-

however, when flownvertically, even withthevortex, it

sions. Reference 21 proposes a maximum bank angle

wasnotpossible forthemodel tomove horizontally past

upset of 7 ° as being the maximum for safe acceptable thevortex.

operations during the final part of the approach. Free- flight results of bank angle perturbations, as a result of

Verticaltrajectories wereflownthrough thevortex

encounters with vortex wakes of various strengths, are

asshown in figures 27and28.Nearly every penetration

shown in figure 30. The data indicate that by using the

resulted in largebankangles andlargelateral displace-

criteria of reference 21, a maximum CLg of approxi-

ments afterencountering thevortex. As a result, nearly

mately 0.95 would be allowable. Note that these tests

all encounters terminated with recovery on the safety

were flown with a control system with relatively high-

cable afterthemodel wentoutof control andexceeded

gain feedback loops that respond more rapidly to upsets thewindtunnel testenvelope.

than a pilot would in a typical business jet aircraft.

Therefore, the comparison with reference 21 criteria is Vortex Encounter Data Trends for illustrative purposes only.

As previously mentioned, the primary objective of Another indicator which will be a key pilot-observed the current test was to determine whether the free-flight response is the roll rate generated by a vortex encounter.

test technique could be used to fly a model in the pres- Figure 31 shows the roll rates, converted to full-scale ence of a wake vortex flow field. Because of the prelimi- values, which would be experienced during encounters nary nature of the test, time constraints prevented an with vortices of various strengths. These values compare exhaustive data set with which to make statistically valid with a maximum roll rate achievable by the airplane, in conclusions; however, the data may be analyzed to show free air, of approximately 33.5°/sec (dampers off). In some general trends.

addition to bank angle upsets, significant lateral displace- ments occur as a result of a vortex encounter. The lateral The vortical flow field produced varying rolling moments on the model, depending not only on the vortex displacements are particularly undesirable on a precision strength, but also on the relative position of the model approach to landing where very accurately flown ground and the wake vortex system. An example of the rolling track paths are required. Figure 32 shows the maximum lateral translation velocities across the ground that moment produced by the vortex on the model, as a per- cent of roll control power available, is shown in figure 29 occurred during the tests. The data show uncommanded lateral velocities up to 20 percent of the approach for a CLg = 1.18. This figure is a combination of all dynamic data available when the model was at approxi- velocity.

mately the 0 ° radial location from the reference vortex position. The data show a slightly increasing right rolling Conclusions moment as the model gets closer to the vortex, with a A wind tunnel study was conducted to determine the maximum occurring at about 50 in. (approximately one feasibility of using the free-flight test technique to study half the span of the following model) from the vortex.

wake vortex encounters. A generic business class jet air- For example, as presented in reference 20, these trends plane model was instrumented and flown in the vicinity are consistent with general rolling moment predictions of a wake vortex generated by a rectangular wing. The for this span ratio. At the maximum, the roll requirement strength of the vortex was varied by adjusting the is approximately 70 percent of the available roll control available. As the model moves nearer to the vortex refer- generating-wing angle of attack and allowed study of a range of generating-airplane pairs (with span ratios of ence position, the control requirement rapidly shifts to a 0.75) and separation distances without the uncertainties left rolling moment due to the vortex, which exceeds the in vortex decay and atmospheric effects. The study roll capability of the model as it approaches the vortex.

showed thatthefree-flight testtechnique wasa viable difficulty of maintaining control. Specific conclusions reached are as follows:

anduseful toolin thestudy of wake vortex encounters--

combining vortexflow fields,airplane dynamics, sen-

1. The free-flight test technique can be used to fly a sors, andflightcontrol aspects.

model in the vicinity of and to conduct encounters with a

Dataobtained duringthis testincluded qualitative

vortex generated from upstream models.

andquantitative results. Thetestindicated thatalthough

2. Accurate model and vortex-positioning data dur-

eachof the longitudinal controlschemes provided an

ing flight can be derived by using postflight photogram-

easy-to-fly airplane, the g-command system reduced the

metry measurement techniques.

longitudinal upsets slightly more than the pitch rate or angle-of-attack command systems. Steady-state limits of 3. Mapping of approximate induced rolling moment controllability during flight were documented as a func- due to the vortex flow field can be conducted.

tion of vortex strength. These data showed limits of con- trollability for steady flight to be at a vortex strength 4. Effects of vortex encounters on airplane model corresponding to a lift coefficient of approximately 0.7 responses can be measured and repeated in well- on the generator wing. By flying several vortex encoun- controlled conditions with known vortex strengths.

ter trajectories at high vortex strengths, a mapping of roll Results showed vertical trajectories to be the most angle, roll rate, lateral velocity, and vortex-induced roll demanding for encounters with high-strength vortices, rate acceleration was conducted. The data quantified the and lateral penetrations were the most demanding at effects of an airplane entering vortex flow fields of vary- lower vortex strengths.

ing strengths, and just as importantly, demonstrated the ability to fly safely and to recover from scale-model wake encounters in a confined test area. Pilot comments NASA Langley Research Center indicated that the selected flight trajectory through the Hampton, VA 23681-2199 vortex and the vortex strength affected the perceived June 12, 1997 Appendix back, removing pitch rate feedback, or adding ct feed- back for static stability augmentation by using either one or both of the wingtip boom vanes. The pitch pilot could Control Laws for Follower Model change these parameters with switches on the pitch con- trol box during a flight.

Basic Considerations The flight control laws for the model were developed Pitch rate command mode. The pitch rate com- to provide good flying qualifies for flying a model in the mand mode was entered by selecting a switch on the wind tunnel. The operational envelope modeled during pitch pilot control box. This command mode combined the tests was the landing-final approach configuration at the pitch rate and pilot command signals through a for- low speed and sea level altitude. The control laws were ward path integrator. Additional proportional feedback implemented on a digital minicomputer using an update of pitch rate was added after the integrator. This mode rate of 200 iterations/see. All flight control system gains changed the pilot input commands to represent pitch rate could be modified in flight by a computer operator. commands, and with centered stick, it was essentially a Selected flight control system switches and gain paths pitch attitude hold system.

could be activated by switches on the pilot control boxes.

Block diagrams of the flight control laws are shown in g-command mode. The g-command mode was figures AI through A4. Switch default values and gain selected by switches on the pitch pilot control box. This values shown in the figures correspond to the baseline mode combined the pilot commands with normal accel- (g-command) system used throughout most of the eration through the forward path integrator. Proportional testing. pitch rate feedback was used outside the integrator. With centered control stick, this mode attempted to maintain lg flight and resulted in a more lively model than the Longitudinal Axis previous modes. Pitch rate damping was increased to There were several control law options in each of the provide better flying qualities.

lateral and longitudinal axes. The longitudinal axes had either proportional or proportional plus integral modes Lateral-Directional Axes using either normal acceleration or pitch rate in the inte- The lateral-directional control laws included an grator. The feedback gains for pitch rate, angle of attack, and normal acceleration could be adjusted in flight. aileron-to-rudder interconnect (ARI) so that one pilot could fly the roll-yaw axes of the airplane with one con- Angle-of-attack command mode. The angle-of- trol. Switches on the roll-yaw pilot control box enabled attack command mode was the most simple control law selection of roll rate, yaw rate, side acceleration, or side- scheme. This mode bypassed the integrator by setting a slip from one or both wingtip booms to be selected as switch to zero. The only parameter in the feedback loop feedbacks during flight. Additionally, differential signals with the _ command system was pitch rate. The pitch of angle of attack between the two wingtip mounted vanes could be used as a feedback.

pilot had the option of adding normal acceleration feed- J N a o ,-2 < ]3 cs_ cs_ c_ o

E_ o

F, Q

J

E_

I I rl

i

E_

_C_

I I i

I

_0

Eo

_ b

i

L_

E_

tm ,Ak References of Hazard Posed by Lift-Generated Wakes. AIAA-93-3518, Aug. 1993.

1. Hinton, David A.: Aircraft Vortex Spacing System (AVOSS) 12. Rossow, Vernon J.: Wake-Vortex Structure From Lift and Conceptual Design. NASA TM- 110184, 1995.

Torque Induced on a Following Wing. AIAA-93-3013, July 1993.

2. Rossow, V. J.: Inviscid Modeling of Aircraft Trailing Vortices.

Wake Vortex Minimization, NASA SP-409, Feb. 1976, 13. Greene, G. C.: An Approximate Model of Vortex Decay in the pp. 9_0.

Atmosphere. J. Aircr., vol. 23, July 1986, pp. 566-573.

3. Bilanin, A. J.; Teske, M. E.; Donaldson, C.; and Snedeker, 14. Somers, D. M.: Design and Experimental Results for a R. S.: Viscous Effects in Aircraft Trailing Vortices. Wake Vor- Flapped Natural-Laminar-Flow Airfoil for General Aviation tex Minimization, NASA SP-409, Feb. 1976, pp. 61-128.

Applications. NASA TP-1865, 198 t.

4. Lee, T.; and Lan, C. E.: Navier-Stokes Calculation of Wing 15. Coe, Paul L., Jr.; Turner, Steven G.; and Owens, D. Bruce: Wake Structure. AIAA-94-1882, June 1994.

Low-Speed Wind-Tunnel Investigation of the Flight Dynamic Characteristics of an Advanced Turboprop Business/ 5. McGowan, W. A.: Calculated Normal Load Factors on Light Commuter Aircraft Configuration. NASA TP-2982, 1990.

Heavy Transport Airplanes. NASA TN D-829, 1961.

16. Ciffone, D. L.; and Orloff, K. U: Application of Laser Veloci- 6. McMillian, O. J.; Schwind, R. G.; Nielsen, J. N.; and metry to Aircraft Wake-Vortex Measurements. Wake Vortex Dillenius, M. F. E.: Rolling Moments in a Trailing Vortex Flow Minimization, NASA SP-409, Feb. 1976, pp. 157-192.

Field. NASA CR-151961, 1977.

17. Riley, Donald R.; Brandon, Jay M.; and Glaab, Louis J.: 7. Stuever, Robert A.; and Greene, George C.: An Analysis of Piloted Simulation Stud), of an ILS Approach of a Twin- Relative Wake-Vortex Hazards for Typical Transport Aircraft.

Pusher Business Commuter Turboprop Aircraft Configuration.

AIAA-94-0810, Jan. 1994.

NASA TM-4516, 1994.

8. Garodz, L. J.; Lawrence, D. M.; and Miller, N. J.: Measure- 18. Childers, Brooks A.; Snow, Walter L.; Jones, Stephen B.; and ment of the Trailing Vortex Systems of Large Transport Air- Franke, John M.: Support of Wake Vortex Detection Research craft, Using Tower Fly-by and Flow Visualization--Summar); in Flight and Wind Tunnel Testing Using Videometric Tech- Comparison and Application. FAA-RD-75-127, Jan. 1976.

niques. Paper presented at ISPRS Commission V Intercongress (Available from DTIC as AD A021305/8.)

Symposium, Mar. 1994.

9. Brashears, M. R.; and Zalay, A. D.: Laser Doppler Velocime- 19. Military Specification: Flying Qualities of Piloted Airplanes.

ter Measurements of B-747 Wake Vortex Characteristics.

MIL-F-8785C, Nov. 1980.

Proceedings of the Aircraft Wake Vortices Conference, 20. Rossow, Vernon J.; and Tinling, Bruce E.: Research J. N. Hallock, ed., FAA, Mar. 1977. (Available from DTIC as on Aircraft/Vortex-Wake Interactions to Determine Accept- AD A055510.)

able Level of Wake Intensity. J. Aircr., vol. 25, June 10. Smith, H. J.: A Flight Test Investigation of the Rolling 1988, pp. 481-492.

Moments Induced on a T-37B Airplane in the Wake of a B-747 21. Sammonds, Robert 1.; Stinnett, Glen W., Jr.; and Larson, Airplane. NASA TM X-56031, 1975.

William E.: Criteria Relating Wake Vortex Encounter Hazard 11. Rossow, V. J.; Sacco, J. N.; Askins, P. A.; Bisbee, L. S.; and to Aircraft Response. J. Aircr., vol. 14, no. 10, Oct. 1977, Smith, S. M.: Measurements in 80- by 120-Foot Wind Tunnel pp. 981-987.

Table I. Geometric and Mass Characteristics of Follower Model Geometric characteristics: Fuselage: Length, ft ............................................................... 7.873 Maximum diameter, in ...................................................... 11.2 Wing: Area (trapezoidal reference), ft 2 ............................................ 9.869 Span, ft ................................................................ 9.072 Quarter-chord sweep, deg ................................................... 1.41 Aspect ratio ............................................................... 8.3 Taper ratio (trapezoidal reference) ............................................ 0.35 Mean aerodynamic chord, in ............................................... 14.172 Dihedral, deg .............................................................. 4.0 Horizontal tail: Area, ft 2 ................................................................ 2.067 Span, ft ................................................................ 3.211 Aspect ratio ............................................................. 4.988 Quarter-chord sweep, deg ................................................... 31.6 Dihedral, deg ............................................................. -3.0 Taper ratio .............................................................. 0.35 Mean geometric chord, in ................................................. 8.324 Vertical tail: Area, ft 2 ................................................................ 2.016 Height, in .............................................................. 18.223 Quarter-chord sweep, deg .................................................. 50.0 Mean geometric chord, in ................................................. 16.259 Mass characteristics: Weight, lb .................................................................. 92.5 Moment of inertia: Ix, slug-ft2 .............................................................. 4.636 I r, slug-ft 2 ............................................................. 14.547 Iz, slug-ft 2 ............................................................. 16.666

Table II. Generating-Wing Vortex Strength

O_g C L g Fins F fs -6.0 -0.929 -8.80 -120.22 -4.0 .054 5.14 70.27 -2.0 .195 18.45 252.08 0.0 .360 34.08 465.46 2.0 .516 48.92 668.24 4.0 .669 63.36 865.45 6.0 .810 76.72 1048.03 8.0 .945 89.53 1222.99 10.0 1.071 101.50 1386.42 12.0 1.185 112.22 1532.91 13.4 1.247 118.13 1613.65 Table III. Dynamic Scaling Relationships [In current test, N = 0.175] Parameter Model Full scale MS Time ....................................

MS FS,_ Linear velocity ............................

Linear acceleration ........................ FS MS FS MS,,_ Angular velocity ..........................

FS MS Angular acceleration .......................

N 3 --xFS Weight ..................................

N 5 Moment of inertia ......................... --xFS (I N _°xMS -xFS Dynamic pressure .........................

N t_

Table IV. TestSetup Geometry

Component Downstream distance, x/bg

Generating-wing quarter-chord Laser 1 0.84 Laser 2 1.92 Laser 3 2.61 Table V. Sensor Accuracies and Resolution Free-flight model sensors Range Accuracy Resolution +200°/sec _+0.7°/sec Roll rate gyro 0.1 °/sec +200°/sec :k0.7°/sec 0.1 °/sec Yaw rate gyro +200°/sec +0.7°/sec Pitch rate gyro 0.1 °/sec A x accelerometer +20g +0.002g 0.01g +_20g Ay accelerometer +0.002g 0.01g A z accelerometer _0g +0.002g 0.01g Boom -10 ° to 80 ° ½2 ° 0.06 ° +30 ° a-L-_l o Boom 13 0.03 ° +30 ° a-L-_l o Aileron position 0.02 ° _+30 ° a-&lO 0.03 ° Elevator position +_30 ° a-&_lO Rudder position 0.03 ° aEstimated based on repeated calibrations during test.

Table VI. Model Dynamic Characteristics Model scale, Full scale, Level I requirements Parameter dampers on/off dampers on/off (ref. 19) 0.24/0.38 0.56/0.89 <1.4 x r, sec .......................

0.98/0.69 2.35/1.65 <1.8 t30, sec ......................

40/80 16.7/33.5 Pss, deg/sec ...................

2O

Tracking cameras

Power and

control cable

Roll

Tunnel ground board Flight control computer Tunnel operator Figure 1. Sketch of free-flight test technique.

L-93-4368

Figure 2. Generating winginLangley 30-by60-Foot Tunnel.

£

2_ _162_i._ _ _0.2_c 14.17

38.53

108.86

. -_ _6-- 3996_

___ -____11,__

Figure 3. Free-flight model. Dimensions are in inches.

C L

I

-5 0 5 10 15

_,deg

Figure 4. Generating-wing lift coefficient.

L-94-01277

Figure 5. Model in vortex during free-flight test.

Figure 6. Wake vortex free-flight testsetup.

i.8 -- q3 [] [] Free flight [] Ref. 14 1.6 -- []o 1.4 -

CL

_o 1.2 - o [] l.O - O o I_ I I I I I I .8 -2 0 2 4 6 8 10 12 _, deg Figure 7. Flight-determined lift characteristics of free-flight model.

2F

Free flight -2 -4 -6 _e, deg -8 -10 -12 I I I I I I -14 0 2 4 8 10 6 12 or, deg Figure 8. Trim tail settings in flight.

-.0025--

-.0026 -.0027 -.0028 CI_ a -.0029 -.0030 , , , I , , , I , , , I , , , I , , , I , , , I , , , I 0 2 4 6 8 10 12 c_, deg .00034 - .00032 .0003O .00028 CI8 r .00026 .00O24 , , , I , , , I , , , I , , , I , , , I , , , I , , i I .00022 -2 0 2 4 6 8 10 12 _, deg Figure 9. Roll control effectiveness (ref. 14).

_--_ /- Generating win_ _pwas h

_, Downwash _ J_

vo ex

Upwash _ _ _-- Jpwash _ _ ,_v rotation Vortex rotation Figure 10. Flow-field schematic.

z, in.

-10 1-10 x, in.

Figure I 1. Photogrammetry-determined vortex position dynamics. Data span 0.5 sec; q = 5.3 psf.

Average roll

control required, 50

percent

o....o---

I I I I I .2 .4 .6 .8 1.0 CLg Figure 12. Roll control required for steady flight in vortex.

8 m 7- 6- 5- 4- 3- I I I I I 0 .2 .4 .6 .8 1.0 CLg Vortex-induced differential angle of attack at wingtip probes while flying in center of vortex core.

Figure 13.

3O

-.08

-.07

-.06 Maximum static roll-cont -.05 ACI -.04 -.03 -.02 -.01 I I I I I 0 .2 .4 .6 .8 1.0 CLg Figure 14. Linear prediction of rolling moment based on induced differential angle of attack at wingtips.

Vortex generating wing

Plan view Rear view

Figure 15.Signconvention formodel location relative tovortex.

Count 800

1.2 1.4 1.6 1.8 2.0 2.2 2.4 Downstream distance, X/bg Figure 16. Histogram of downstream distance between model and generating wing. All runs included.

0v=270 °

R

0v = 180 °

0 V ----- 0 ° Approximate location of port vortex 0 v = 90 ° Figure 17. Right-to-left horizontal encounter. CLg --- 0.95.

Ov I I I I I I I I I I I Distance R,in.

I

or 180

I

0v,deg

I I / I/ II tl

, I f

.10

.08

.06

.04

.02

Clv 0

-.02

-.04 L

-.06

-.08

-.10

_a _e

Control

positions, 0

/

deg -5

-10

-15

-20

,,_--I..,.m

-25

Time Figure 17. Continued.

P q r 2O

: A A

Angular rates,

deg/sec

-20 -40 -60 !

-80 Attitude, -i0 deg B -20 -30 B --40 i I -50

/

& Flow angles at left wingtip J / probe, deg I ii ,,--\ I

"[. k/

-5 I i -10 i i

;/

! I -15 0 4 6 8 10 Time Figure 17. Concluded.

0 v = 270 ° R 0 v = 0 ° 0 v = 180 ° Approximate location of port vortex 0 v = 90 ° Figure 18. Left-to-right horizontal encounter. CL_ = 0.95.

R 0v

L

Distance R,in.

or 180 0v, deg

J

i J J i I i J I I I I i CI v _a _Se Control positions, deg /l_'/ 1%-%

L_

t I | l I I I I 2 3 Time Figure 18. Continued.

I

Angular rates, L

deg/sec 0

-20

-40

-60

-80 : , , , , I I I I I I | !

lo:_ .........................

o

deg -10 : -20 p.

p- -30 -40 " -50 i ....

sill 15 a p Y Flow angles at left wingtip 0 probe, deg _,_ x. I \_ -5 _* t .I -10 f I I t I I I I I I I a I 1 s I I -15 0 4 Time Figure 18. Concluded.

0v =270 °

R 0 v = 180 ° e V _- 0 ° Approximate location of port vortex 0 v = 90 ° Figure 19. Right-to-left horizontal encounter. CLg = 1.07.

0v L R I I 300 I I I I I Distance R,in.

I \ or 0v, deg I I I II|l C1v

/

I I I I IIII _a 8e Control positions, deg till litl i t i t t t 1 2 4 _me Figure 19. Continued.

4O q p r Angular rates,

deg/sec

-20 -40 -6O -80 ! I I I I I I I 1 I ! I I I I I I I I I 3O (D 2O IO f Attitude, -10 deg -20 -30 -40 -50 I I i i I I I I I I ! ! 1 | I 1 [ I t 1 IO Flow angles ,,... J -\ \ atleftwingtip probe, deg ! \ t \ \ A -5 -10 I 1 1 I I t ! !

-15 2 3 4 Time Figure 19. Concluded.

0v =270 °

R

Start

0v= 180 °

0 V _ 0 ° Approximate location of port vortex 0 v = 90 ° Figure 20. Descending vertical encounter. CLg = 1.07.

R Ov Distance R,in.

or 180 0 v, deg 0 I ! I 1 I | l | I I 1 !

.10 .08 .06 .04 .02 \f CIv -.02 X.,,l --.04 -.06 -.08 -.10 I I 1 !

L 8a 8e Control positions,

'k_

deg -5 -10 -15

j

-20 : -25 il|l ! 1 I I I I I I I111 2 3 4 5 Time Figure 20. Continued.

/

Angular rates, deg/sec 0 ,___. F

_7

-20 ,._ -40 -60 -80 .... i ! | i

l

E___ I f/ Attitude, 7"-, -10 deg -20 -30 -40 -50

L

i0 Flow angles at lefiwingtip

°_

probe, deg X. _/'\ -5 -10 t i | | I ! I | -15 ....

0 2 3 4 5 Time Figure 20. Concluded.

0v=270 °

R

8O 0 v = 180 ° 0 v = 0 ° Approximate location of Start port vortex 0 v = 90 ° Figure 21. Ascending vertical encounter. CL, = 1.07.

R 0v f Distance R, in.

/ or 180 / 0v, deg / | i t ! I t IIII llll ii It .10 .08 .06 .04 .02 Civ 0 -.02 --.04 -.06 -.08 I11[ IIII I I I I I I I I -.10 _a _Se

l

Control positions, 0 deg -5 -10 -15 -20 : n l t i IIII II |l i s i i IIII -25 4 5 Time Figure 21. Continued.

Angular rates, deg/sec 0 -20 -40 -60 -80 ¢ tg Attitude, deg -10 -20 -30 -40 III I -50 Flow angles N at left wingtip 0 \ probe, deg \ -5 \ \ -10 I I ! I ! I I I I I i i -15 ....

0 3 4 Time Figure 21. Concluded.

0v= 270 °

R

0 V = 0 ° 0 v = 180 ° Start / Approximate location of port vortex 0 v = 90 ° Figure 22. Right-to-left horizontal encounter. CL_ = 1.18.

R 0v Distance R, in.

or 180 0v, deg / I l r-- -r --J-- "i- i | | | 1"" .10 .08 .06 .04 .02 CIv 0

\

-.02 -.04 -.06 -.08 -.10 |111 iii1 _a

25 f

L 8e

Control positions, 0 deg -5 !

-10 1 I -15 \1 -20

L___

iiii -25 I I l I till 0 2 3 Time Figure 22. Continued.

r\

40--

Angular rates,

deg/sec 0

\

-20

-40

-60

Illl 1 I i I I

-80

4,.. ° Attitude,

deg -10

-20

/

-30

-40

III1 III I

-50

If'

\

Flowangles

/

atleftwingtip 0

probe, deg

-5

/

-10

I [ I | I I 1 l I t I I

-15

0 2

Time Figure 22. Concluded.

5O

0v =270 °

R

2O

0v= 180 °

O v= 0°

L Approximate location of port vortex 0 v = 90 ° Figure 23. Left-to-right horizontal encounter. CLg = 1.18.

R 0v Distance R,in.

, ,,,,,,,

or 180

0v,deg

\

120 ,.%

\ \ !

IIIl Ill| .10 .08 .06 .04 ..

.02

CIv

r L, 0 _xNf f_'A -.j -.02 :_- V

\

-.04 : -.06 .

-.08 1 i I I III l -.10 l , l * 25 : _a _Se 10..

Control

5__ /h

.g positions, A oVV -v A',.

deg

- F -5 -lo : k/

V

-15 _ -20 E -25 _" ' ' ' ' Time Figure 23. Continued.

P r Angular rates,

deg/sec 0

-20

-40

-60

_1111 iiii Illl

-80

: : 0 Attitude, deg -10 -20 -30 -40 IIII -50 Ot

i ,.q . /

_J _" \ Flow angles I :'7_, at left wingtip 0 I probe, deg

S

I -5 I -10 It -15 .... I 0 2 Time Figure 23. Concluded.

0v= 270 °

R

I00

0 v = 180 ° 0 v ---- 0 ° L Approximate location of port vortex 0 v = 90 ° Figure 24. Descending vertical encounter. CLg -- 1.18.

R 0v \ \ Distance R, in.

\ or 180 \ 0v, deg i i | t I | I ! i i i i I i 1 | | i i i .10, .08 .06 .04 .02 Clv 0

\#

-.02 -.04 -.06 -.08 i ! i i -.10 1 I i 1 25 : _a 2O _e 10 : Control positions, 0 deg -5 \_ f -10 - ,q -15 -20 -25 I I I I I I I I I I ! I !

Time Figure 24. Continued.

/ .

Angular rates, deg/sec 0 -20 -40 -60 I [ ! l -80

(p

.I" ..r_ 7 Attitude, -10 deg -20 -30 -40 Illl Illl -50 ....

(_

L

Flow angles at left wingtip 0 probe, deg -5 -10 ! l i I -15 ....

0 5 Time Figure 24. Concluded.

0v= 270 °

R

0v= 180 °

0 ev =0°

Approximate

location of

\

port vortex

0 v = 90 ° Figure 25. Ascending vertical encounter. CLs = 1.18.

R Ov / Distance R,in.

/

or 180

/

0v,deg

/ t

/

! l 1 1 .10 .08 .06 " .04 .02 O.

CI v

7.

-.04 -.06 -.08 1 i i ! I | | • 1 i i ! ! ! I t i i -.10 : , , , ,

21F

Control

positions, • _ .t 0 _ ,/_,\

deg

-15 _ -I_I I i i l I I | i 2 3 4 5 Time Figure 25. Continued.

Angular rates, deg/sec -20 -40 -60 -80 oo / Attitude, -10 deg IIII Flow angles at left wingtip 0 probe, deg \ IIII 3 4 5 Figure 25. Concluded.

0v =270 °

R

6O

0 v = 0 °

0v= 180 °

Approximate location of port vortex 0 v = 90 ° Figure 26. Right-to-left horizontal encounter. CLg = 1.25.

R 0v Distance R,in.

or

0v,deg

0 lilt

i i .1o .08 " .06 .04 .02 _ 1

CI v

V v : V -.02 -.04 : -.06 -.08 -.10 • , , , I 2O

Control

positions,

deg

-5

V

-10 -15 -20 -25 II1,| IIII II11 IIII iii 0 2 3 4 Time Figure 26. Continued.

P

I

Angular rates, 0 _X _ - ----_,4 deg/sec --" -20 -40 -60 III| lea| -80 ....

f X, 0 "_ " •_ jJ Attitude, deg -10 -20 -30 .

-40 II*L i a I i -50 ....

P

I \ | Flow angles ,...,

I

atle_wingtip 0

. V

probe, deg I -5 rJ | \J I I -10 i i a I llal -15 ....

2 3 4 5 Time Figure 26. Concluded.

0 v = 270 ° R Start 0 v = 180 ° 0 V = 0 ° Approximate location of port vortex 0 v = 90 ° Figure 27. Descending vertical encounter. CLg -- 1.25.

R 0v \ Distance R,in.

OF 180 I 0v, deg

f"

I I I I .10 .08 .06 .04 .02 CI v -.02 -.04

F

-.06 -.08 -.10 _a 20 _e I Control positions, deg -5 -10 -15 -20 -25 I I I I I I I I IIII IIII Time Figure 27. Continued.

P q r 40 .......................

,.J

20.

Angular rates, deg/sec -20 -40 -60 -80 ll|l |lit till I I I I ¢

t

.f J J -...._./ Attitude, -10 deg -20 -30 -40 i0 Flow angles at left wingtip probe, deg -5 d / -10 IIII I I I I 1111 Illl -15 , , , A 3 4 Time Figure 27. Concluded.

0 v = 270 ° R 0 V = 0 ° 0 v = 180 ° Approximat_ location of port vortex Start 0 v = 90 ° Figure 28. Ascending vertical encounter. CLg = 1.25.

R I t 0v It Ixi Distance R,in.

or 180

I

0v,deg

I I I

0 _llll II]l

i i i s i i i

.10

.08

.06

.04

.02

Clv 0

-.02 "-_ -.04 _,_ /" -.06 .... "'- -.08 -.10 ....

I I i i i | i I i | | i _a 2O 6e

L

7.

Z Control positions, deg f -5 /_\ /IV '_ \ -10 "_ ". " \ / -, / -15 -20 -25 , , , i i 1 i i i i i t i 3 4 5 Time Figure 28. Continued.

P "1 r Angular rates,

deg/sec

\

-20

-40

-60

t ! s |

-80

¢ Attitude,

-10

deg

-50 ....

f Flow angles I ".,_ at le_wingtip 0 probe, deg -5 \

J

\ / -10 --15 I I I I I I I * | | .I I.

0 1 2 Tim e Figure 28. Concluded.

Maximum right 100-

CIv,

percent

-50

Maximum left -100

, , I q j , I ; , t I , ; , I , , , I _ ; , I , , , I -1500 20 40 60 80 100 120 140 Distance, in.

Figure 29.

Rolling moment produced on model due to vortex interactions. CLe = 1.18 along 0 ° radial +30 °.

0 -- O -10 -20 Maximum roll angle, deg -30 O0 _ 0 0 -40 -50 I I I I I I I 0 .2 .4 .6 .8 1.0 1.2 1.4 CLg Figure 30. Maximum roll angle produced by vortex encounter.

-5 -10 Maximum full-scale -15 roll rate, deg/sec -20 -25 I -30 I I I I 0 .2 .4 .6 .8 1.0 1.2 1.4 CLg Figure 31. Maximum roll rate produced by vortex encounter.

20- Lateral translation, 0 0 percent 10 forward speed ,, , I, ,, I, ,, I,, ,1 , , ,I, ,, I _ ,, I 0 .2 .4 .6 .8 1.0 1.2 1.4 CLg Figure 32. Maximum lateral translation velocity produced by vortex encounter.

7O REPORT DOCUMENTATION PAGE Fo_ A_o_oved OMB No. 0704-0188 Public repoding burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of thh collec'_on of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Joffersol Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503_ 1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 13. REPORT TYPE AND DATES COVERED November 1997 ] Technical Paper 4. TITLE AND SU,',/i/LE 5. FUNDING NUMBERS Application of Wind Tunnel Free-Flight Technique for Wake Vortex Encounters WU 505-64-13-15 8. AUTHOR(S) Jay M. Brandon, Frank L. Jordan, Jr., Catherine W. Buttrill, and Robert A.

Stuever 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESSEES) 8. PERFORMING ORGANIZATION REPORT NUMBER NASA Langley Research Center L-17462 Hampton, VA 23681-2199 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORINGffAONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA TP-3672 Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES Buttrill: Unisys Corp., Hampton, VA.

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE Unclassified-Unlimited Subject Category 08 Availability: NASA CASI (301) 621-0390 13. ABSTRACT (Maximum 200 words) A wind tunnel investigation was conducted in the Langley 30- by 60-Foot Tunnel to assess the free-flight test tech- nique as a tool in research on wake vortex encounters. A typical 17.5-percent scale business-class jet airplane model was flown behind a stationary wing mounted in the forward portion of the wind tunnel test section. The span ratio (model span-generating wingspan) was 0.75. The wing angle of attack could be adjusted to produce a vortex of desired strength. The test airplane model was successfully flown in the vortex and through the vortex for a range of vortex strengths. Data obtained included the model airplane body axis accelerations, angular rates, attitudes, and control positions as a function of vortex strength and relative position. Pilot comments and video records were also recorded during the vortex encounters.

14. SUBJECT TERMS !15. NUMBER OF PAGES Free flight; Wake vortex; Wind tunnel 16. PRICE CODE A04 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF REPORT OF THIS PAGE OF ABSTRACT OF ABSTRACT Unclassified Unclassified Unclassified NSN 7540-0%280-5500 Standard Form 298 (Rev. 2-89) Prescribed by ANSI Std. Z39-18 298-102