Paraglider Recovery System for the Saturn Booster

8/7/2019 Paraglider Recovery System for the Saturn Booster

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The development of the Flexible Wing represents a major advance in the field of aerodynamic structure,providing an extremely lightweight, aerodynamic lifting surface. Ryan engineering evaluation has lead to variousapplications fo r the wing.

The Report presents the results of a technical and economic feasibility study to evaluate dr y land recoveryof the Saturn booster (employing the Rogallo Flexible Wing), and is submitted in accordance with the requirementsof NASA Contract NAS-8-I50I. Both the C-l and C-2 Saturn booster configurations were analyzed at the request ofNASA with major emphasis directed toward the C-2. The investigations may be considered applicable to any otherbooster configuration, falling within the spectral burn-out conditions fo r the C-l and C-2.

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Basic wing aerodynamic data fo r the performance and stability and control investigations were derivedfrom wind tunnel tests conducted by Langley Research Center and Ryan. The figure shows estimated lift/drag ratioand maximum lift coefficients fo r both rigid leading edge and inflatable leading edge booster recovery systems usedin the study analysis. Uncorrected data of the expected booster configuration obtained from a Ryan conducted windtunnel test is also shown. Corrections to this substantiate the estimated data used in the performance study.

The higher the lift/drag ratio, the more favorable the gliding range. The rigid leading edge configurationis superior aerodynamically because of lower sweep angles obtainable by spreader bar. Inflatable v.i ngs, wherespreader bars are not used tend to seek higher sweep angles and, therefore, reduce lift/drag ratios.

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~ / E S T I M A T E D RIGID 1........,.1' 1 l ~ , t i I I R I I . UNCORRECTEDEOOE MODEL

,,' '

tI LIFT

ESTIMATED INFl.ATABl.f: DRAG.- L A D l N G ~ ' ! " : > ~ ' . . V----.L.-.____ ------- ---

o _10 20 30 40 50-ANGLE OF ATTACK - a - DEGREE

LIFT DRAG RATIO vs ANGLEOFATJACK" , )-; ..,', '

CL-MAX. RIGID= 1.15CL MAX .INFLATABLE=.95


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Studies of the booster burnout cover velocity ranges of 2000 to 3000 ft. per sec. at low altitudes fo r theC-2, and 8000 to 8500 ft/sec at high altitudes for the C-l. Trajectory studies of these burnout conditions indicate thatthe C-2 recovery requires a high dynamic pressure deployment capability, while the C-l deployment will occur atvery low dynamic pressure. Due to higher energy of C-l missions, re-entry will occur at a considerable distancedown range from launch site. Dry landing fo r these missions must occur at an offshore island installation. TheC-2 missions, in contrast, may be studied for return flight to launch site. Booster missions within this burnoutvelocity-altitude-spectrum should yield conclusions similar to those studied.

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III 280,000I 240,000I -. C-l.....I '7'C) 200,000->. .

PACKAGING &EJECTION

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The suggested nose gear and rear main gear uses conventional landing skis. The nose gear, mounted onthe forward booster and containing a tungston contact surface for minimum drag, is extended on command by apneumatic system. The aft main gear is also retractable and provides a copper contact surface for maximum drag.Utilizing these two different contact surfaces will give a degree of directional stability to the booster during groundrun. I f this directional stability is marginal, use of a nose wheel in place of the sk i is proposed. The rear mainand outrigger extension schematic is also shown.

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STA. 781.3 (REF.)I

24.0 STROKE

TUNGSTEN BonOM

. -1---50.50 .. I

NOSE GEAR

STA. 189.5 (REF.) STA. 118.0 (REF.)I

FmlNG INCLUDESBOOSTER TIEDOWN LUG

15.0 STROKE OlEOS LATCH INOOWN PosmON

~ ~ ~ ~ 1--------- 107.50----\-----1COPPER BOTTOMAFT MAINPRESSURE BOTTLE

PRESSUREAIR MOTOR REGULATOR PRESSURE REGULATOR

OLEO AIR MOTOR

OLEO 1lE0

REAR MAIN AND OUTRIGGER SCHEMATICLANDING GEAR

OLEO

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The outrigger skis are used to provide lateral support to the booster during the ground run. A tungstoncontact surface is provided fo r min. drag. The system is retractable and stored within the booster container.

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STA. 118.0 (REF)I B. L. 88.25 (REF) LOCALLY DISHED TO ACCOMODATE GEAR" -_I EXPLOSIVE BOLT

o 1-- / ------- _ ------------- ... '" I,-- ------I ---- --___. --_.-------. --- ~ ' _ r .._------ . ----- - ' \

- - - - - - - - _ - / ,I . ~ " ~ ' -.... ..r----i. . ' ~ /_/ \!I

DOWN POSITION LATCH

AIR MOTOR

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TUNGSTEN BOTTOM.\ .\1---2---60.00 -----1

OUTRIGGER SKI SYSTEM

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Shown here is the control cabling which provides c. g. shift for the turning and pitch maneuvers. Threeseparate hydraulic motors ar e required, all driven by a single hot gas motor. A single hydraulic motor providespitch control cable actuation. The remaining two hydraulic motors, when actuated differentially, provide c. g. shiftfo r the turning maneuver, and, when driven in unison and in conjunction with the pitch control hydraulic motor, pro-vide booster attitude orientation in the flight path.

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The control power unit suggested fo r the system considers the conventional hot gas motor with hydraulicdriven power motors. Fuel is provided by the helium pressurized hydrazine tank.

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PITCH CONTROL SPROCKET CHAINHOT GAS MOTOR

HYDRAZITE TANK

HELIUM TANKROLL CONlROL CABLES ~ = - : : : : : : : ~ ~ : : ; j

ROLL CONlROL CABLE

ROLL CONTROL WINCH 2 REQ'D. - ~ ~ ~ ~ ~ ~ ~ - ~ ~ = ~ ~ ~ ~ ~ - - ~ - 4 . . J L - ~ - -: : : :JROLL CONTROL HYD. MOTORS 2 REQ'D. _______ CENTER BEAtAPRESSURE REGULATOR HOT GAS MOTORHYD. FILTER

GAS GENERATORTHROTTlE VALVE:

/SERVO VALVE/EXHAUST BY PASS VALVE

HYD. PUMP

' . ' -- - - - - - - - - - ~ \.LEADING EDGE BEAM

PITCH CONT. CABLEROLLER CHAIN

CONTROL SYSTEM SELF PRESSURIZING RESERVOIR

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The static longitudinal stability of the wing booster system is illustrated here by the wing position angle {3as a function of wing angle of attack O!. Three separate c. g. positions ar e shown: the booster maximum forwardand af t with residual fuel, and the neutral or zero fuel case. The area of negative or diving moments is shown inthe left side of any particular c. g. trim line, while the right side is presented by positive or upsetting moments. Asseen from the figure, the system is completely stable from an angle of attack of 30{3 or lower. Any disturbance,such as a gust which brings momentary angle of attack change from trim position, will be compensated by the properrestoring moments. Between 30 and 40 degrees, one degree of neutral stability exists, and near CL maximumlongitudinal stability again is reached. The separation of each of the c. g. trim lines illustrates the high tolerancerequirement placed on knowing the c. g. position if a satisfactory glide angle (and therefore angle of attack) is required.

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Requirements fo r stability augmentation fo r the C-2 recovery system are illustrated in this figure. Theydo not, however, apply to al l flexible wing recovery systems. Different missions and applications of the wing callfo r different stability requirements.

1. Dynamic analysis indicated a slight phougoid during glide mode, though subsequent analysis leavessome doubt of it s existance. This phougoid is believed detrimental to the landing maneuver because of the longperiod oscillation about the flight path.

2. The c. g. boundary limits experienced by any reSidual fuel will be sufficient to offset the desiredflight path i f c. g. is not known and proper corrections made. A system is required, therefore, to provide the correctangle of attack in the flight path regardless of center of gravity location.

3. A pitch rate equivalent to 1. 5 ft/sec. is incurred on landing i f booster attitude is not controlled.Although not sufficient reason in itself for stability augmentation, this is one of the advantages to be gained by useof such a system.

4. Wind tunnel data analysis indicates that the C-2 recovery system possesses a marginal degree oflateral stability at high angles of attack. To accomplish dry land recovery, turn maneuvers at high lefts ar e required.

5. Characteristic of al l high altitude vehicles is the decrease in wing damping at high altitudes. Lowlateral stability at high angles of attack require stability augmentation i f high lift turns are necessary.

6. Spiral divergence, characteristic of al l winged vehicles, must be eliminated.

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1. PHOUGOID --

2. CG TRAVEL -----

3. PITCH RATE. EQUIVALENTTO 1.5FT./SEC. ON LANDING

4. DIFFICULTY AT FLYING HIGH LIFT TURNS------5. INSUFFICIENT DAMPING AT HIGH ALTITUDE

AT HIGH LIFT

6. SPIRAL DIVERGENCE _ - --=_::::-.i

REQUIREMENTS FOR STABILITY AUGMENTATION

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A typical landing profile for the recovery system as a function of altitude and horizontal approach andlanding distance is given in this figure. Approach speed is 110 lmots, with 60 feet per second sink speed. The boosteris oriented to the ground horizontal attitude by the flight control reference system. Flare is initiated at approxi-mately 250 ft . altitude. Flare is controlled automatically from the ground by a programmed rate of sink as afunction of altitude. Booster touchdown occurs between 60 and 70 lmots and 2 to 5 ft . per second sink speeds.Booster ground contact attitude is 5 or less. The wing is allowed to stay attached fo r approximately 280 feetground run distances to provide booster directional stability during ground run. Thereafter (and at a velocity ofapproximately 40 lmots) the wing is released from the booster and falls back to the ground. The booster totallanding ground ru n distance is approximately 653 feet.

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500

L.&.I 300t:::c t

200

100

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: ~ . HO KNOTS.!60 FT./SEC .'0/ .

100 KNOTS .f FT ./SEC90 KNOTSf ? r l20 Fl./SEC .

T.D.6070 KNOTS25 Fl./ SEC . 40 KNOTSr:= - 553 Fl:-r - - 2 Fl; 1 .. ' ... j

O ~ ~ ~ ~ ~ ~ ~ - - ~ - - ~ - - ~ ~ ~ ~ ~ ~ ~ o 200 400 600 800 1,000 1,400 1.800 2,200 2,600 3,000 3,200 3,400 3,600HORIZONTAL DISTANCE

LANDING PROFILE

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Landing ground run distance fo r 30 ft. per second head wind, tail wind, and no wind conditions ar eshown on this figure as a function of the ground friction coefficient. The copper contact surface for the main gearlanding skis provides the greatest retardation, and distances, therefore, can be estimated using this value. Check-ing with the Figure, it is seen that ground run distances of over 1,000 feet ar e not expected.

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0.8

.....0.6c..,)......

......L.I.JC)c..,)::z:

30 FT.jSEC.HEAD WIND ZERO WINO 30 FT./SEC . TAIL WINO

52 COPPER0 . 4 t - - - - - ~ r - - - - - - - - - - l , . _ _a::......

TUNGSTEN0.2 t - - - - - - _ : _ _ _ -

o 200 400 600 800 1,000LANDING GROUND - RUN

FRICTION COEFFICIENT vs LAND ING GROUND-RUN

1,200 1,400

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Various operational stages of the booster reuse cycle for cost evaluation are shown here.

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TRANSPORTTO STORAGE

TRANSPORT FROM FAB.TO STORAGE STORAGE

TRANSPORT TO REFURBISHINGSITE

LAUNCH

===-=::;::==n1

TURN &GLIDE

BOOSTER RE-USE CYCLE

DEPLOYMENT

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The probability of booster reuse in a program of 12 firings a year is shown as a function o.f the number ofboosters required. Probability values are based on the accumulative reliability of the various segments andcomponents of the total recovery system and it s mission. Analysis indicates that maximum booster reuse probabilityis .732, while minimum probability reuse is .579. These values represent expected fabrication and refurbishingrequirements to sustain a program of 12 firings per year.

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An additional expression of the probability of reuse is the number of launches per booster required fora given program. For this program of 12 firings a year, the launches per booster varied from 2.38 fo r a probabilityof reuse of .579, to 3.74 launches. This boundary of launches per booster is used in the economic evaluation study.

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III a::: 4.0LA.J

V)I 00QQa:::I LA.J 3.00- ANTICIPATED RANGE OFV)LA.J LAUNCHES PER BOOSTERI:::z=c..;)z::;:,4 : 2.0I -- '

I \ \ ~ ' t FARNICA TIOI(~ \ \ ~ \ ~ ~ \ ~ \ \ t ~ NATEI 1.02 4 6 8 10 12I BOOSTER PER YEAR

I LAUNCHES PER BOOSTER vs BOOSTER FABRICATIONI AND REFURBISHING RATESI


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Comparative cost study results for three different percent refurbishing rates ar e shown here as a functionof average launches per booster vs . booster program savings. The refurbishing rate is expressed as a percentof the original booster cost. A refurbishing rate of 20% is expected for the recovery program because of its auto-matic features and simplicity of the landing maneuver. Cost estimates used in this evaluation are:

Recovery system developmentRecovery operating equipment costs

Storage facilityLaunch check equipmentDeploy-glide and Landing

Landing facilitiesFlight & landing control equip.

Inspection and Refurbishing facilitiesBooster transportation equipment

Personnel requirements per yearRecovery system unit costBooster unit cost

147,000300,000

10,000,0001,350,0003,000,000

500,000

15,000,000

3,300,000375,000

9,000,000

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4.5CI:La.I1 -.V )==: 3.5L.r.ICL..V ).....:z: 2.5S.....c,.s:,cc 1.5cc

PROGRAM COST WITHOUT .RECOVERY =1.3 BILLIONPROGRAM DURATION 12 YRS.LAUNCH RATE 12 YRS.

E=60 E=40 E= 20_.-..... ....._.-..-._-----_. _._----------------- ---

100 200 300 400 500 600 700BOOSTER PROGRAM SAVINGS (MILLIONS)

AVERAGE LAUNCHES PER BOOSTER vs BOOSTER PROGRAM SAVINGS

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Additional savings obtained from use of the recovery system are represented by percent in boosterdollars per pound of payload in orbit VS number of launches per booster. As shown, considerable savings accrueregardless of rate of refurbishment.

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I 45I t::::al 40cc:I C):z: 35-cI ex:C) 30-- '>-ex:I CL. 25'-'-C)I al 20 E-0.60-- ''-t:A- 15I :z:-V ') 10c..:::s - MOST PROBABLE RANGE .-I :z:-:>ex: 5V ')I 0 0 I1.5 2.0 2.5 3.0 3.5 4.0 4.5I LAUNCHES PER BOOSTERI PERCENT SAVINGS OF PAYLOAD COST WITH ARECOVERYI SYSTEM vs LAUNCHES PER BOOSTERI

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For other missions, the recovery system weight will increase only slightly. It should also be rememberedthat the sink speeds of 5 ft . per second or less are possible fo r any recovery system.

The conclusions of the booster recovery program ar e based primarily on the C-2 missions.

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PHA E IANAL YSIS & DEVElOPMENTENGINEERINGWINO TUNNEl MODELS1/ 6 SCALE MODElSTABILITY CHUTE TESTSFULL SCALE GROUND DEPLOYMOCK UP

e i&;IJANALYSIS & DETAIL DESIGNFABRICATIONSYSTEM STATIC &ENVIRONMENTAL TESTSFLIGHT TESTSGROUND SUPPORT EQUIPMENT

.6- - PHASE I

DEVELOPMENT PROGRAMPHASE /I -

DESIGN FREEZE +

' M&& ..

- . - - .6

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CONCLUSIONS1. C - 2 DRY LOAD RECOVERIES ARE POSSIBILITIES2. PACKAGING WITHIN C - 2 COUNTERS IS POSSIBLE3. PACKAGE "TYPE" RECOVERY SYSTEM INSTALLATION IS POSSIBLE4. RECOVERY SYSTEM WEIGHT EQUALS 8% OF RECOVERED WEIGHT5. SINK SPEEDS EQUAL TO 5FT./SEC. OR LESS ARE POSSIBLE

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Paraglider Recovery System for the Saturn Booster