evaluation of separation mechanism design for the … · evaluation of separation mechanism design...

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National Aeronautics and Space Administration

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Evaluation of Separation Mechanism Designfor the Orion/Ares Launch Vehicle

Abstract:As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed 2-D (Simulink) model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the 2-D model vs. a full 3-D (ADAMS) model are discussed as well.

https://ntrs.nasa.gov/search.jsp?R=20100042391 2018-05-28T05:28:50+00:00Z

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Evaluation of Separation Mechanism Design for theOrion/Ares Launch Vehicle

Kevin E. Konno, Daniel A. Catalano, Thomas M. KrivanekNASA John H. Glenn Research Center, Cleveland, Ohio

39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight Center, May 8, 2008

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Service Module

Crew Module

Solar Arrays (2)

Service Module (SM)

Engine Nozzle

ARES 1 Stack

Orion vehicle (fairings removed)

Launch Abort System

Orion

Upper Stage

1st Stage

Hardware Overview

Spacecraft Adaptor (SA)

Service Module (SM)

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2

1

1st stage recovery

4

3

5

67

2nd stage disposal

8 9 10 11

MET(sec) EVENT

1 0 12 stage Ign2 58.6 Max q3 125.3 1st stage sep4 126.3 US ign5 137.9 Begin pitch profile6 153.3 SA Fairing Jett7 156.3 LAS Jett8 590.5 MECO9 621 CEV Sep10 1568 Orb insert11 1740 SAW Deploy

1

1st stage recovery

3

2

4

5

6

2nd stage disposal

78 9

MET(sec) EVENT

1 0 1st stage Ignition23

125.3 1st stage sep

4126.3 US ign

56

153.3 SA Fairing Jett

7

156.3 LAS Jett

8

590.5 MECO

9

621 CEV Orion Sep1568 Orb insert1740 SAW Deploy

Nominal Ascent Timeline

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Nominal Separation Event

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Separation mechanisms traded:

Compression springs Pyrotechnic Gas Thruster Pneumatic actuators

Separation Spring

Frangible Nut

Preload Rod

Bolt Catcher

•Low part count•High reliability•Well known•simple

•Higher specific thrust (~10x springs’)

•Higher part count, possibly lower reliability

•Higher specific thrust (~5x springs’)

•Higher part count, possibly lower reliability

Graphics supplied courtesy of Scot, Inc.

From 1983 to 2005, Spacecraft and Fairing separation systems accounted for 10% of all launch failures, according to AAS 03-071 paper. Vehicle dynamics accounted for another 4%.

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Run vehicle separation simulation for abort case to

size actuators

Run simulations for max/min engine thrust

Run simulations for max/min dump rate

Run simulations for dispersions on mass properties

Design parameters traded/optimized:• Actuator type• Actuator force/stiffness• Actuator stroke• Spacecraft Adaptor (SA) flange diameter

Obtain nominal

vehicle Mass, Moment of

Inertias, dimensions &

State Conditions

Assume actuator separation system force/spring rate,

travel, & number of actuators

Min. Clearance Achieved?

Output clearance, time, force, and distance

Yes

No3-D(ADAMS) and 2-D (Simulink) models(Dynacon and ADAMS for LMCO)

Design study variables: • Upper Stage residual thrust (0 - max lbs)

• Vehicle dump rate & direction (0-35 deg/sec)

• Spring-out condition (1 in 12)

• Vehicle mass property dispersions (+/- 10%)

Preliminary Design Process

Update System mass, reliability

Models correlate-NASA/NASA? NASA/LMCO?

Yes

No

Investigate discrepancies

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0

500

1000

1500

2000

2500

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5

time, sec

Forc

e, k

g

Spring Force

Gas Thruster

1095 J (808 ft-lbf)

1144 J(844 ft-lbf)

Total impulse for Spring is 1095 J vs Gas thruster @ 1144 JAverage accel for Spring is 0.026 g's greaterActuator induced velocity for Spring is 0.477m/s vs 0.485 m/s

Both achieve clearance at 4.83 sec

Characteristics of Force Application

Spring Actuator

Gas Thruster Actuator

Simulink results

Pneumatic or

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0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

2500 4500 6500 8500 10500 12500spring stiffness, kg/m

clea

ranc

e,

m

3.08m (121.6") flange dia.- array

3.82m (150.4") flange dia.- nozzle







allo

wab

le

SA Flange Size Optimization (considering arrays and nozzle clearance)

Arrays will clear the SA for any flange diameter smaller than 145”. Engine nozzle will clear SA for any flange diameter larger than 125”.

ADAMS results

J-2x Engine on, Dump rate fixed at 5 deg/sec

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2-D model vs 3-D (ADAMS) model comparison

0

0.2

0.4

0.6

0.8

1

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5

time, sec

Sepa

ratio

n C

lear

ance

, m

ADAMS Global - Nozzle Clr

ADAMS Global - Solar Array Clr

Simulink Global - Nozzle Clr

0

0.1

0.2

0.3

0.4

0.5

3500 5000 6500 8000 9500 11000 12500 14000

actuator spring stiffness, kg/m

clea

ranc

e, m

AdamsSimulink

0.051m Clearance Required

Nozzle clearance vstime curves for a typical design case

Nozzle clearance vsspring stiffness curves for a typical design case

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“Natural SeparationNatural Separation” Concept

For 2 bodies attached and under constant pitch rate, for a given protrusion diameter and length, upon separation:•There exists a cavity diameter D that the protrusion will naturally clear at, regardless of pitch rate. •The bodies will separate and protrusion will clear body 2 at a prescribed angular rotation regardless of pitch rate

This neglects outside forces acting on the bodies, which can easily be considered later in design process

Body 2

Body 1Body 1

protrusion

Cavity diameter

Pitch rate

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Making use of Natural SeparationNatural SeparationFalcon 1 Demo Launch- Staging anomaly

Camera mounted here looking down

Cavity diameter

Protrusion length

*All information borrowed from SpaceX public website: www.spacex.comFalcon 1 stack

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Orion Natural SeparationNatural Separation Dynamics Benefits At Work

-0.1

-0.05

0

0.05

0.1

0.15

0.2

10 15 20 25 30 35vehicle dump rate, deg/s

clea

ranc

e, m

Con

tact

Posi

tive

Cle

arac

ne3.4m (135") flange dia., No actuation or J-2X

.051 m Required Clearance

4,893 kg/m 8,018 kg/m 6,250 kg/m

10,358 kg/m

3.4m (135") flange dia., No actuation w/J-2X

3.15m (124") flange dia., No actuation w/J-2X

added spring stiffness to meet req'd clearance

Vehicle pitch rate, deg/sec

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Recontact at Low Dump Rate

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• For two bodies joined and tumbling at a constant angular velocity, when separated will each maintain that same angular velocity after separation (neglecting outside forces)

• For spacecraft mechanism design sizing, the abort/off-nominal case is not always the driving design case

• Independent analysis and verification of critical vehicle dynamics can be beneficial in avoiding costly corrections later

• Intelligent preliminary sizing of spacecraft separation mechanism geometry sensitive to separation dynamics can improve overall mission reliability and save on mechanism weight, especially if Natural Separation concepts are invoked early the design

Lessons Learned

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Conclusions

• Lower fidelity, 2-D equations of motion model can be very useful in separation mechanism design. It provides insight into separation events and the many parameters and their relative sensitivities.

• A more detailed 3-D geometric dynamics model is helpful in considering out of plane effects which may be significant such as CG offsets, single actuator/spring failures, and product of inertia terms.

• For the Orion crewed vehicle separation system a simple mechanical spring mechanism has been chosen as the baseline design because the spacecraft geometry was sized efficiently, minimizing the required actuator force even with significant force margin (25%) applied.

AcknowledgmentsThe authors would like to acknowledge the contributions, advice, and suggestions of Keith

Schlagel and Lance Lininger of Lockheed Martin Corporation which aided in the development and compilation of this work.

Further information: Restricted NASA TM Spacecraft Separation System Dynamics for the Orion/Ares launch Vehicle. To include vehicle mass properties, full Simulink code, tank slosh modeling.

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Backup slides

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0

1

2

3

4

5

6

0 10 20 30 40 50 60

Vehicle Dump Rate, deg/s

Tim

e to

Min

imum

Cle

aran

ce,

sec

0.051m Clear Constant ForceNo actuator forceActuator, constant forceNo Actuator

F=1,567 lbs

F=1,148 lbs

In all no-actuator cases spacecraft rotates through 28o to achieve separation

F=0 lbs

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Simulink Flowchart

Calculate Moment Arm & Vehicle Cg’s

Vehicle Mass, Moment of Inertias, dimensions & State

Conditions

Input separation system spring rate,

travel, & number of springs

Input RCS System force, number of engines, timing

Input J-2X TVC angle & residual

thrust levelVehicle initial rotation rate

J-2X Engine nozzle length, diameter &

Orion pedestal diameters & height

Min. Clearance Achieved?


Iterate to time step

Calculate translational & rotational velocity from

equations of motion & forces

Integrate velocities

Yes

Yes

No

NoDetermine CLV and Orion rotation & relative position


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