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SPACE LAUNCH SYSTEM

National Aeronautics and Space Administration

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Modeling and Test of Space Launch System Core Stage Thrust Vector Control

Jeb S. Orr, Ph.D. / Draper (Jacobs ESSSA) MSFC Control Systems Design and Analysis Branch (EV41)

George C. Marshall Space Flight Center

.1

Aerospace Control and Guidance Systems Committee Meeting 116 March 15-18, 2016

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Introduction

.2

•  Space Launch System (SLS) – NASA-developed launch vehicle for large-scale

(exploration-class) crew and cargo access –  Shuttle-derived hardware and processes leveraging

Constellation program development experience (tanks, engines, boosters)

–  Primary development configurations are 70t crew (Block I) and 130t cargo (Block II)

•  SLS Thrust Vector Control (TVC) Actuators –  SLS uses a total of 12 TVC DoF (boost phase) and 8 TVC

DoF (core phase) –  TVC performance is critical for stability, loads, and

integrated vehicle control – A novel approach to analysis and test has been

undertaken to verify and validate TVC models used for flight dynamics and control design

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Heritage TVC System Considerations

.3

•  SLS TVC actuators are Shuttle heritage – Quad-redundant, mechanical feedback hydraulic actuator – Closed-circuit hydraulic power provided by redundant APUs

• GHe (core stage), hydrazine (booster) – Robust dynamic pressure feedback (DPF) provides active load

damping over a wide range of load resonances – Core stage structure, interfaces, hydraulic support system, and TVC

Actuator Controller (TAC) are a new design – There exists a need to update and certify existing high-fidelity

models prior to flight

SLS combines a novel modeling approach with preflight testing to anchor model predictions

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Modeling Methods

.4

Prescribed Motion Engine in FEM and locked. Actuator coupled to global vehicle model. Load approximated by spring. Ghost modes are a problem. F&V method.

Standard Model Engine in FEM and locked. Rigid engine in system EoM. Ghost modes are a problem. ASAT & FRACTAL 1 method.

Reduced Body Model Engines removed from FEM. Load approximated by spring. Good approximation for global vehicle dynamics. Ghost modes eliminated. FRACTAL 2 method.

Coupled TVC-FEM Engines and springs removed from Simplex. TVC-servo dynamics coupled to local FEM. Higher fidelity for local dynamics and coupling effects. Multiple engines. MASV method.

Traditional Methods

New Methods

Loads Model Actuator approximated by spring. All FEM. Cannot model servodynamics. Overconservative for load resonance (0.5% damping).

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u  The STS SSME TVC actuator is robust to load resonance variations within the Orbiter design range

•  The single-spring load resonance frequency is given by

where are the nozzle angular and total linear system stiffness, is the actuator moment arm, and is the engine inertia

u  The servoactuator DPF network phase stabilizes the load resonance (active damping) u  Analysis shows sensitivity to values outside of the Orbiter load frequency range

•  Stability of the actuator (inner loop) is affected – linearization of DPF may not be accurate •  SLS FCS uses advanced servoelastic feedback model to aid in global bending stabilization

Motivation for Detailed Modeling

5

! =

s(Kn +KTR2)

Jn

Kn, KT RJn

Simplex TVC Open Loop

phase (deg)

gain

(dB)

−900 −720 −540 −360 −180 0

−50

−30

−10

10

30

50

70 Orbiter Type III nominal (8.6 Hz)Orbiter lo (6.5 Hz)Orbiter hi (10.8 Hz)

28.4 dB4.77 Hz

80.5°2.70 Hz

108.6°6.95 Hz

44.5°21.11 Hz

101−45

−40

−35

−30

−25

−20

−15

−10

frequency (Hz)

gain

(dB)

bode magnitude, axis = 2, time = 51

Low Stiffness Load

High Stiffness Load Sensitivity in global vehicle structural

dynamics

SLS Autopilot Open Loop Response (typ.) Typical Shuttle Orbiter Inner Loop

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TVC Model V&V Using MASV

.6

GR FRT Test Profile MASV

TVC Complex Model

ER35 Lab Testing

GR FRT Test Profile

Green Run FRT

MASV (Test Correlated)

TVC Simplex Model

STE Models (GR FEM)

VM Tools (FRACTAL) Flight

STE Models (Flight FEM)

•  Multiple Actuator Stage Vectoring (MASV) Model – Developed by Draper to improve modeling of interactions between TVC

servodynamics and local structure –  Reduce risk and increase understanding of core stage TVC dynamics –  Verify TVC performance and stability using high-fidelity structural response

•  Eliminate single-spring approximation of load compliance –  Used along with “Complex” single-axis model and 2-axis ILS (lab testing) to verify TVC FRT test procedure (excitation and data recovery)

– MASV validated using GR FRT data and used to parameterize VM Simplex model (prediction)

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Multiple Actuator Stage Vectoring (MASV) Model

.7

•  Approach –  Engine dynamics are replaced with a high-fidelity modal representation of

the core stage thrust structure – Allows coupling of multiple actuators with a single set of dynamic modes – A partitioning procedure is used to identify and group generalized

coordinates that do not contribute to dynamic response to reduce the number of DoF

 

Actuator    

R  

β  

z  

x  2 2,f Φ

 1 1,f Φ

 

Gimbal  

Rocket  body  

-630 -540 -450 -360 -270 -180 -90 0-40

-30

-20

-10

0

10

20

30

40

Phase [deg]

Gai

n [d

B]

Open Loop Nichols Plot

MASV: RigidMASV: Rigid + 840 Modes

Shuttle Orbiter Type III High Bandwidth Verification Case

Modal model recovers test-correlated spring approximation response

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Frequency Response Testing

.8

•  FRT is necessary to characterize TVC behavior in flight-like boundary conditions

–  Space Shuttle Orbiter used a dedicated test article (MPTA) and an extensive test program to reduce TVC modeling uncertainty • 12 static firings from 1978-1981

•  SLS will execute a limited test on flight hardware at the Core Stage Green Run (GR)

– Determine frequency response and transient response of the coupled actuator-structure system in hot-fire conditions

–  120 second test window at 109% PL –  Instrumented using existing flight piston position

TM sensors and drag-on string potentiometers –  Testing reproduces boundary conditions and

effects that are difficult to model & predict accurately, especially coupling, gimbal friction, oil air entrainment, thermal drift, etc.

TABLE 1.1 - TDS VERSUS TEST HATRIX..... --, _

1RSD _ 4120/01 JJ4ATRIX STATIC FIRING (Codo on last page)

Test Number Tank "l-S 2-5 3-S 4-S 5_-F. 5-F 6-F 65-, 75 8S ] 95 10F iIF--'-I'IS-_ 12F _Duration(sac) 1.0 20 42 104 1.5 54 19 555 554 540! 574! 550 550 586 625

TRSO Paragraph ,Jhrust(J)..... i0 70 _ _p I00 100100 100 1001_001102 102 i02 I_OOand Gimb°l _- ., Yesye; YosYes(Ye_4YesYe, Yes"Yes Yes

Test Requlr_ent POGO Yes Yes Yes Yes Yes YesJ YesJ YesEngine Out 90 4854855055_0_ 15_01402429 4_ _5

" TDS 5.1.1 TVC Performance Evaluation _ ." " , J J_ I. Slnusoldal Engine Glmballln_ ........... I I_ - 0.2" /_nplltude , . P I Iel J F I -_ - 0.4 degrees amplItude ,:' " =Pj =Pj e .,. I .,I zFsl.tF=I F' I F. - 0.6 degrpes amplitude P FI

2. Step Response ;-_ : .... IP/ IPI . P. IFI IFII FI'J Fi. 3. Ramp Response . .... IPI IPI P . , [FI F Fm F,- 4. Strokln_ Response ....... F. 5. Flight Profile F,-- 6, Simultaneous Engine Glmballln_/Throttlln_ .... F . •,. 7. Routine Engl,ne Positioning/S SHE Side Load Evaluatlor P P P P ' F'--m | _ ,_

8. Engine, Clearance Checks (Non-Firing) P F' II 9. Hydraulic Power Fal'lure Simulation (flon-Flrln_) P [Fi F

10. TVC Channel Failure Simulation F

,- I1. Halting Ramp Response , IPI .... m ..... IF/, IF) , F m12. Slmultep.eous Glmballln_ annd POGO r,. F

:TOS 5. l.2 Hydraulic System Performance verification .....(Note: Engine Glmbellln 9 Requirements are defined ......

.... In TDS 5.$.1) ........_ I. Evaluate Hydraulic System during Engine .P IPI IP! P [ P P F ....._ Throttling - , ,,_ 2. Evaluate Hydraulic System durlp9 Steady State P P P : e IFI -P P P F I :

Engine Performance (no Throttling) ............ 3. Hydraulic Wormant,Flow Evalu,atlon ....

Constant Return Pressure P P P P P P P P F

nstantSuppI emperetu;. i i I P I P I PI FI I I I- KSCCountdownTlmellne ...... I I I I |PI I P I (F!I ' ' ' F'I ' :1 "1 I

mRequlrements for stub nozzles have been met.Requirements with flight nozzles are still open.

Instrumentation locations on base heat shield

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u  FRT profile is executed in the thrust vector null space of the CSEs •  Profile results in no net commanded off-axial loads on the stage structure •  Some small loads will result due to non-ideal tracking of the commands, stage

structural dynamics/asymmetry, actuator/engine variability •  Commanded in two channels (null pitch, null yaw) @ 50 Hz, 120 sec, 109% PL •  Low-frequency and high-frequency ID on each engine on orthogonal DoF •  Transient ID (varying amplitude step response) on each channel

FRT Profile Design

9

PROFILE CHANNEL 1 PROFILE CHANNEL 2

140 160 180 200 220 240 260 280−1

−0.5

0

0.5

1

firing time (sec)

engi

ne c

omm

and

(deg

)

Profile Channel 1

140 160 180 200 220 240 260 280−1

−0.5

0

0.5

1

firing time (sec)

engi

ne c

omm

and

(deg

)

Profile Channel 2

LF ID

LF ID

HF ID

HF ID Transient

Transient

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150 155 160 165 170 175 180 185−0.5

0

0.5

firing time (sec)

engi

ne a

ngle

(deg

)

commandresponse

150 155 160 165 170 175 180 185−1

−0.5

0

0.5

1

firing time (sec)

pist

on p

ositi

on (i

n)

u  All maneuvers are individual sinusoids with start-stop buffers of 3 settling periods • Minimum of 3 periods or 8 setting times, whichever is longer •  Enables frequency domain recovery using least squares, much more accurate than FFT with

sine sweep in noise environment if command profile is known • Multisine cannot be easily mechanized with null constraint and system is not linear

u  Low frequency ID maneuver consists of 8 sample-aligned frequencies (log spacing) • Reach full command amplitude (quarter-period alignment) @ 0.4 deg Z-T-P (STS MPTA) •  There are no sample-aligned frequencies between 6.25 and 12.5 Hz @ 50 Hz rate

Low Frequency ID

10

Predicted response (no noise)

Concurrent testing on coupled axes is possible through frequency separation

since single-component frequency-domain LSQ is used for signal recovery

Channel 1 (Hz) Channel 2 (Hz)

0.40-6.25 Hz increasing

7.0-14.0 Hz increasing

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190 195 200 205 210

−0.5

0

0.5

firing time (sec)

engi

ne a

ngle

(deg

)

commandresponse

190 195 200 205 210−1

−0.5

0

0.5

1

firing time (sec)

pist

on p

ositi

on (i

n)

u  High frequency ID maneuver consists of 8 non-sample-aligned frequencies •  Log spacing from 7 Hz-14 Hz (bounds predicted nominal

closed-loop load frequencies with ~25%-30% margin) • Command amplitude increased to 0.8 deg Z-T-P to

increase SNR on piston measurement

High Frequency ID

11

Predicted response (no noise)

Channel 1 (Hz) Channel 2 (Hz)

7.0-14.0 Hz increasing

0.40-6.25 Hz increasing

194 195 196 197−1

−0.50

0.5

firing time (sec)

engi

ne a

ngle

(deg

)

194 195 196 197−1

0

1

firing time (sec)

pist

on p

ositi

on (i

n)

Transient buffer Integration time

Sample alignment effect

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215 220 225 230 235 240

−0.5

0

0.5

firing time (sec)

engi

ne a

ngle

(deg

)

commandresponse

215 220 225 230 235 240−1

−0.5

0

0.5

1

firing time (sec)

pist

on p

ositi

on (i

n)

u  Transient ID maneuver consists of 3 positive and negative steps at 0.2, 0.4, and 0.6 degree amplitude •  Similar procedure to STS; Opposite channel is quiescent during step •  6 settling times between steps (~2 seconds) and 2.5 second persistence time •  Evaluate cross-axis coupling, load effects, push-pull symmetry, amplitude nonlinearity,

bias, scale factor error, drift •  Limited resolution/quantization/noise can limit utility of steps at very small amplitudes

Transient ID

12

Predicted response (no noise)

-O.& • [ .......,. I - _.. , ,,

I_ " J ,...... •'C.G. 11_KING •U ["0,8 ....

..... 1, .J

TV_ AClURTOi_"'_ -I., - ' _'

POSITION, ....... _.

INQIES _ °

' . ,.. .,,I

p..,a, J_

-- Sllll SIIPARATION _/- ...... _.:

_'__1 - FIR,_r 30 SE _CONllS " _o_D_

.,., l I °leo t;_O 14e 16_ 180 _ee 2_0 24e _6e 28e 3ee 3_e 9nJ_

Tlr_ IN S(¢ONDS OR(F"TIM£1 ie,eoo.oo._loeT.s?e)eqlgF0'_5"m

O_ ,-.t-0

T Figure 2.3.9-5 "rvc Actuator PoaitJ.on Versus Time for SSNE 2 Yaw Durino Flioht Profile _ _OJ ilil)_lllilicj Test 0 =Ln , "I

STS MPTA data (position, in)

233.5 234 234.5 235 235.5

−0.10

0.10.2

pist

on p

ositi

on (i

n)

Simulated position data with noise (0.2 deg step)

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u  Frequency-domain reconstruction using a describing function-like approach • Given an unknown SIS(M)O nonlinear system described by

with a known input and stochastic noise n, an estimate of the linear frequency response (first harmonic, dependent on amplitude A) is computed from using the Fourier coefficients (k=number of integration periods)

•  Implemented in discrete time using 50 Hz trapezoidal integration. • Correction for ZOH delay is applied to post-processed complex arrays.

Data Processing

13

ż = f(z, u)

q = h(z, u) + n

u = A sin!t

a01 =2

kT

Z kT/2

�kT/2q(t) cos(!t) dt

b01 =2

kT

Z kT/2

�kT/2q(t) sin(!t) dt.

|N | =pa021 + b

021

A\N = tan�1

✓a01b01

◆

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u  Good frequency ID of engine position and load resonance is possible with noise and quantization error

Frequency ID Results

14

10−2 10−1 100 101 102 103−80

−60

−40

−20

0

frequency (Hz)

gain

(dB)

Piston position test points

10−2 10−1 100 101 102 103−400

−300

−200

−100

0

frequency (Hz)ph

ase

(deg

)

140 160 180 200 220 240 260 280−1

−0.5

0

0.5

1

firing time (sec)

engi

ne a

ngle

(deg

)

commandresponse

140 160 180 200 220 240 260 280−1

−0.5

0

0.5

1

firing time (sec)

pist

on p

ositi

on (i

n)

10−2 10−1 100 101 102−80

−60

−40

−20

0

frequency (Hz)

gain

(dB)

Engine position test points

10−2 10−1 100 101 102−400

−300

−200

−100

0

frequency (Hz)

phas

e (d

eg)

140 160 180 200 220 240 260 280−1

−0.5

0

0.5

1

firing time (sec)en

gine

ang

le (d

eg)

commandresponse

140 160 180 200 220 240 260 280−1

−0.5

0

0.5

1

firing time (sec)

pist

on p

ositi

on (i

n)

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u  Test profile verification on the MSFC SSME TVC Inertial Load Stand

Lab Testing

15

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u  The SLS Program has leveraged a unique combination of advanced analysis techniques and testing to validate TVC models for flight

u  Flight control stability and performance is assured with high confidence based on extensive flight experience with high performance NASA heritage hydraulic actuators

u  Test and performance data collected throughout this effort will directly support flight certification as well as post-flight reconstruction and anomaly resolution

Summary

16