April 2011 

NASA/TM-2011-217083 

 

Recent Experiences of the 

NASA Engineering and Safety Center (NESC) 

Guidance Navigation and Control (GN&C) 

Technical Discipline Team (TDT) 

Cornelius J. Dennehy/NESC 

Langley Research Center, Hampton, Virginia 
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April 2011 

NASA/TM-2011-217083 

 

Recent Experiences of the 

NASA Engineering and Safety Center (NESC) 

Guidance Navigation and Control (GN&C) 

Technical Discipline Team (TDT) 

Cornelius J. Dennehy/NESC 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

Acknowledgments 

The author thanks the following individuals: the members of the NASA 

Engineering and Safety Center Guidance Navigation and Control (GN&C) 

Technical Discipline Team (TDT) for their willing and dedicated contributions of 

their collective energies towards the goal of improving the GN&C discipline 

within NASA; the other members of the NESC assessment teams with whom the 

GN&C team collaborated; and the NESC Leadership Team at NASA’s Langley 

Research Center for their support and encouragement in the work of the GN&C 

TDT in general and the development of this paper in particular. The author would 

also like to extend his appreciation to Scott Glubke (Goddard Space Flight Center, 

Code 590) for his rigorous review of and commentary on this paper. 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration. 
iii 

Table of Contents 

1.0 Introduction ................................................................................................................. 1

2.0 NESC Vision, Mission, and Organization ................................................................... 2

3.0 NESC GN&C TDT...................................................................................................... 4

4.0 Experiences of the NESC GN&C TDT ....................................................................... 5

4.1 Max Launch Abort System (MLAS) Flight Test Assessment ..................................... 6

4.2 Constellation System Preliminary Design Review (PDR) Assessment .................... 10

4.3 Orion Ascent Abort One (AA-1) Loads Assessment ................................................ 13

4.4 Orion CM Boilerplate Water Drop Test Assessment Support .................................. 16

4.5 Assessment of Orion Slosh Impact on Ares-I FCS Stability ..................................... 20

4.6 Kepler Recurring Anomalies ..................................................................................... 30

5.0 Conclusions ............................................................................................................... 32

6.0 References ................................................................................................................. 33

List of Figures 

Figure 1. MLAS FTV. .................................................................................................................. 7

Figure 2. Launch of MLAS FTV from Wallops Flight Facility (8 July 2009). ............................. 8

Figure 3. MLAS CM altitude versus range. ................................................................................... 8

Figure 4. CM altitude versus time.................................................................................................. 9

Figure 5. Comparison of PA-1 and AA-1 flight test configurations............................................ 15

Figure 6. Orion CM boilerplate water drop test at Aberdeen Proving Ground. ......................... 17

Figure 7. Inertial sensor instrumentation mounted inside CM test article. .................................. 18

Figure 8. Representative raw X-axis CG#1 accelerometer data from CM water drop test. ........ 19
iv 

Figure 9. Low pass filtered X-axis CG#1 accelerometer data from CM water drop test. ........... 20

Figure 10. Ares-I launch vehicle shortly after liftoff. .................................................................. 21

Figure 11. High-level Ares-I FCS architecture. ........................................................................... 23

Figure 12. Notional depiction of Ares-I FCS gain and phase margins (first-stage example). ..... 25

Figure 13. Notional depiction of Ares-I FCS gain and phase margins (upper-stage example). .. 25

Figure 14. Ares-I control, flexible body, and slosh mode frequencies vs. time. ......................... 27

Figure 15. Orion upstream NTO tank slosh displacement (0.1% damping; TD71 mission). ...... 29

Figure 16. Kepler spacecraft and photometer. ............................................................................. 30

List of Tables 

Table 1. Ares-I FCS stability margin requirements. ..................................................................... 24

Table 2. Space Shuttle launch vehicle FCS stability margin requirements. ................................. 24
v 

Nomenclature 

AA Ascent Abort 

ACM Attitude Control Motor 

AM Abort Motor 

ATB Abort Test Booster 

CAS Credibility Assessment Scale 

CEV Crew Exploration Vehicle 

CG Center of Gravity 

CM Crew Module 

CxP Constellation Program 

DOF Degree of Freedom 

DSN Deep Space Network 

E2E End-to-End 

ESMD Exploration Systems Mission Directorate 

FCS Flight Control System 

FOT Flight Operations Team 

FTV Flight Test Vehicle 

GN&C Guidance Navigation and Control 

GPS Global Positioning System 

IMU Inertial Measurement Units 

IRMA Integrated Risk Management Assessment 

IRT Independent Review Team 

ISS International Space Station 

JSC Johnson Space Center 

LAS Launch Abort System 

LAV Launch Abort Vehicle 

LC Launch cycle 

LIS Lost-in-Space 

LPF Low Pass Filter 

MAVERIC Marshall Aerospace VEhicle Representation In “C” 

Max-Q Maximum Dynamic Pressure 

MLAS Max Launch Abort System 

MMH Monomethylhydrazine 

MSFC Marshal Space Flight Center 

NASA National Aeronautics and Space Administration 

NESC NASA Engineering and Safety Center 

NRB NESC Review Board 

NTO Nitrogen Tetroxide 

OS Objective System 

PA Pad Abort 

PDR Preliminary Design Review 

PID Proportional-Integral-Derivative 

PRA Probabilistic Risk Analysis 

RID Review Item Discrepancies 

RoCS Roll Control system 
vi 

RRT RID Review Team 

SBD Smart Buyer Design 

SM Service Module 

SR Separation Ring 

TDT Technical Discipline Team 

TVC Thrust Vector Control 
Abstract 

The NASA Engineering and Safety Center (NESC), initially formed in 

2003, is an independently funded NASA Program whose dedicated team 

of technical experts provides objective engineering and safety 

assessments of critical, high risk projects. NESC’s strength is rooted in 

the diverse perspectives and broad knowledge base that add value to its 

products, affording customers a responsive, alternate path for assessing 

and preventing technical problems while protecting vital human and 

national resources. The Guidance Navigation and Control (GN&C) 

Technical Discipline Team (TDT) is one of fifteen such disciplinefocused teams within the NESC organization. The TDT membership is 

composed of GN&C specialists from across NASA and its partner 

organizations in other government agencies, industry, national 

laboratories, and universities. This paper will briefly define the vision, 

mission, and purpose of the NESC organization. The role of the GN&C 

TDT will then be described in detail along with an overview of how this 

team operates and engages in its objective engineering and safety 

assessments of critical NASA projects. This paper will then describe 

selected recent experiences, over the period 2007 to present, of the 

GN&C TDT in which they directly performed or supported a wide 

variety of NESC assessments and consultations. 

1.0 Introduction 

The National Aeronautics and Space Administration (NASA) Engineering Safety Center 

(NESC), initially formed in 2003 in the wake of the Columbia tragedy, is an example of a OneNASA Program. NESC is an independently funded NASA program whose dedicated team of 

technical experts coordinates and conducts objective engineering and safety assessments of 

critical, high-risk projects. The NESC is a strong technical resource for customers and 

stakeholders seeking responsive service for solving the Agency’s difficult problems. NESC’s 

strength is rooted in the diverse perspectives and broad knowledge base that add value to its 

products, affording customers a responsive, alternate path for assessing and preventing technical 

problems while protecting vital human and national resources. NESC provides timely technical 

positions to its customers and stakeholders based on independent test and analysis, not opinion. 

By encouraging alternative viewpoints and ensuring objective reporting methods, NESC is able 

to serve as a uniquely unbiased assessment resource. NESC delivers technical evaluation and 

consultation products in the form of written reports that include solution-driven, preventative, 

and corrective recommendations. In July of 2011, the NESC initiated its 370th technical 

assessment. The NESC communicates its lessons learned from each assessment to NASA’s 

leadership through bi-annual briefings and to engineers through both the Agency Lesson Learned 

system and a series of NESC Technical Bulletins issued periodically. These communication 

channels function to inform the NASA technical community and, therefore, NESC’s customers 

and stakeholders. NESC’s range of services includes testing, analysis, and data review in fifteen 

engineering disciplines. NESC also engages in proactive discipline advancing activities. 
2 

The Guidance, Navigation, and Control (GN&C) Technical Discipline Team (TDT), the primary 

subject of this paper, is one of fifteen (15) such discipline-focused teams within the NESC 

organization. The NASA Technical Fellow for GN&C maintains and leads the TDT. The TDT 

membership is composed of senior GN&C engineers from across NASA’s Field Centers as well 

as from its partner organizations in other government Agencies, industry, national laboratories, 

and universities. 

This paper will briefly define the vision, mission, and purpose of the NESC organization. It 

contains a detailed description of the role of the GN&C TDT and an overview of how this team 

operates and engages in its objective engineering and safety assessments of critical NASA 

projects. Descriptions of several recent, 2007 to present, experiences in which the GN&C TDT 

performed or supported a wide variety of NESC assessments and consultations are also included 

in this paper. 

2.0 NESC Vision, Mission, and Organization 

One of the tenets of an effective safety philosophy is to provide an avenue for independent 

assessment of the technical aspects and risks of critical systems. NESC offers this alternate 

reporting path for all NASA programs and projects. 

The vision that NESC has for itself is to serve as the independent and objective deep technical 

resource of choice for NASA Programs and other government Agencies. As its fundamental 

mission, the NESC strives to set the example for engineering and technical excellence within 

NASA. The primary purpose of this independent and objective organization is to increase safety 

through engineering excellence. NESC collaborates with its customers and stakeholders to 

ensure the safety and success of their programs and projects. A resource for the Agency, the 

NESC is a unique and valuable asset for the high-risk programs that NASA undertakes. 

At the core of the NESC is an established knowledge base of technical specialists pulled from the 

ten NASA Centers and from a group of partner organizations external to the Agency. This ready 

group of engineering experts is organized into 15 TDTs. TDT members are drawn from NASA, 

industry, academia, and other government Agencies. By drawing on the minds of leading 

engineers from across the country, the NESC consistently solves technical problems, deepens its 

knowledge base, strengthens its technical capabilities, and broadens its perspectives, thereby 

further executing its commitment to engineering excellence. 

Maintaining a diverse and broad base of knowledge, keeping informed and engaged with each 

Center and the Agency’s major programs, responding efficiently to requests for assistance, and 

retaining a high degree of independence are the basis of the organizational structure of the 

NESC. There are some 50+ full-time NESC-badged employees, the majority of which are based 

at NESC Headquarters located at NASA’s Langley Research Center in Hampton, Virginia. 

NESC employs over 550 other engineers nationwide (part-time) as the members of the 15 TDTs. 

To achieve the goals stated above, the NESC is organized into six distinct offices: 

NASA Technical Fellows assemble, maintain and provide leadership for the TDTs and are 

stewards for their disciplines. The Technical Fellows serve as the senior technical experts for the 

Agency in support of the Office of the Chief Engineer and the NESC. They are an independent 
3 

resource to the Agency and industry to resolve complex issues in their respective discipline 

areas. While they all lead their own NESC TDTs, some Technical Fellows may also lead 

Agency-wide discipline Working Groups. Specifically, the Technical Fellows are responsible 

for: 1) fostering consistency of Agency-level standards and specifications, 2) promoting 

discipline stewardship through workshops, conferences and discipline advancing activities, and 

3) ensuring identification and incorporation of lessons learned into Agency processes. 

NESC Chief Engineers provide insight into their Centers’ programs and help to coordinate the 

facilities and resources of each Center when required to support NESC activities. NESC also 

proactively exploits its network of Center-based Chief Engineers for outreach to and 

communications with the broad NASA community. The Chief Engineers also coordinate with 

the NASA Technical Fellows in the process of identifying potential discipline issues and 

problems to be addressed proactively by the NESC. 

Principal Engineers use TDT members provided by the NASA Technical Fellows and resources 

arranged by the NESC Chief Engineers to lead independent technical reviews, assessments, tests, 

and analyses. 

The Systems Engineering Office dispositions requests as they come in, performs proactive 

trending analysis and problem identification, and provides other integration and system 

engineering support. 

The Management and Technical Support Office is the business arm of the NESC, taking care 

of the contracting, budgeting, and management of the NESC’s infrastructure. 

Under the leadership of the Director’s Office, these five components come together to form the 

heart of the NESC — the NESC Review Board (NRB). The life cycle, from initial assessment 

plans to interim status briefings to final reports, of every formal activity performed by the NESC 

requires approval of the NRB. The NRB is a vital peer review process for the NESC, which 

directly supports the development of high quality end products for stakeholders. All NESC 

reports must be reviewed and approved by the NRB prior to out-briefing the stakeholders. The 

NRB brings a diversity of thought to the decision-making process. It is an amalgam of people 

representing different Centers, programs, and engineering backgrounds. 

After an activity performed by the NESC has concluded, results are delivered to the stakeholders 

in the form of written engineering reports that include solution-driven preventative and 

corrective recommendations. The NESC strives to set the example for the Agency by providing 

full and appropriate documentation of every activity. Along with each report, lessons learned are 

communicated to Agency leadership and to engineers through avenues such as the Agency’s 

Lessons Learned system, the reports themselves, and the periodic NESC Technical Bulletins. 

In addition to acting on requests from outside of the NESC, another important function of the 

NESC is to engage in proactive investigations to identify and address potential concerns before 

they become major problems. To further this goal, the NESC is currently leading NASA’s efforts 

for independent data mining and trend analysis. The NESC has established a Data Mining and 

Trending Working Group that includes representatives from all NASA Centers as well as 

external to the Agency. 
4 

3.0 NESC GN&C TDT 

The GN&C TDT is a technical resource that supports the NESC and the NRB-approved 

independent assessment teams. The primary purpose of the GN&C TDT is to engage in the 

resolution of GN&C related issues throughout the agency when directed by the NRB or by 

NESC senior leadership. A secondary purpose of the GN&C TDT is to identify proactively 

Agency-wide GN&C engineering discipline issues and problems. 

The NASA Technical Fellow for GN&C is assembles, maintains, and manages the CN&C TDT. 

The GN&C resources (subject matter experts, tools, and test facilities) required to support the 

assessment teams and other GN&C-specific NESC activities come from the TDT. The GN&C 

TDT is cognizant of all GN&C related assessments to ensure adequate and timely GN&C 

expertise support. This is accomplished via bi-weekly teleconference meetings and with annual 

face-to-face meetings of the TDT. These and other communication mechanisms (e.g., a NESCinternal GN&C TDT secure website to post team news and other information) are used to unite 

the TDT members located across NASA. 

The GN&C TDT consists of individuals that are experts in a wide range of GN&C subdisciplines including GN&C systems, GN&C analysis, GN&C components and hardware 

systems (sensors, actuators, interfacing hardware systems), GN&C software, flight dynamics, 

mission design, flight operations, launch vehicle flight mechanics analyses, and launch vehicle 

guidance systems. As mentioned above, this team of experts collectively serves as discipline 

“think tank” to identify potential GN&C issues and problems to address proactively by the 

NESC. 

Given the wide-breadth and depth required to staff the GN&C TDT adequately as well as to 

support multiple assessments simultaneously, a staffing model has been developed to recruit and 

staff the GN&C discipline TDT. This staffing model requires skill sets representing discipline 

systems experts, sub-discipline specific experts, and technical team support personnel. The 

GN&C TDT consists of a “core” group of approximately 20–30 discipline systems experts. It 

also consists of an extended team of about 5–6 specific experts from each of the sub-discipline 

areas of expertise that encompass the broad scope of the GN&C discipline at NASA. These subdiscipline experts are on call-up to the NASA Technical Fellow and to the core team. 

Approximately 100 GN&C experts, the majority of them being NASA Civil Service employees 

from across the Agency, currently comprise the entire NESC GN&C TDT. When the operational 

function of the GN&C discipline TDT is constrained by limited Agency in-house staffing 

resources, additional GN&C discipline expertise from outside the NASA community (e.g., 

industry and academia) are exploited to augment the TDT membership. 

The members of the “core” group are senior level individuals from across the Agency that have 

broad, but expert knowledge. These senior experts have in-depth knowledge of one, or several, 

GN&C expertise areas, but probably not all the GN&C areas of expertise. The individuals who 

make up the TDT’s “core” group possess exemplary leadership and teamwork skills since they 

both represent their Center’s GN&C engineering organization and also serve as the GN&C 

leadership interface to the NESC’s assessment teams. 

The sub-discipline specific experts are individuals that have in-depth experience and expertise in 

a specific GN&C area. The TDT core group defines these specific areas. For example, on the 
5 

GN&C TDT, there will be sub-discipline experts in the following areas: inertial sensors, GPS 

navigation, spacecraft attitude determination and control, stellar/celestial sensors, formation 

flying, flight dynamics, aeronautical vehicle flight control, inter-planetary navigation, flight 

mechanics, reaction wheels, control moment gyros, controls structures interaction, mission 

design, launch vehicle guidance and control, etc. 

The technical support group is the third and last major component of the GN&C TDT. The 

technical support group is a small (about 3–5 people) contingent of individuals that support the 

NASA Technical Fellow for GN&C in the day-to-day management and operation of the GN&C 

TDT. These are typically GN&C engineers with perhaps 5–8 years of professional work 

experience. They contribute routine administrative and technical support (e.g., recording 

teleconference meeting minutes, providing logistics for the annual face-to-face meeting, updating 

the TDT’s internal website, etc.) while at the same time benefiting from the mentoring 

experience of working with the other TDT members. The technical support group, by virtue of 

their role on the TDT, has exposure to a wide range of GN&C problems from across NASA as 

well as the opportunity to witness first-hand the problem solving skills of some the Agency’s 

senior GN&C engineers. This has turned out to be a win-win situation that both benefit the 

operation of the TDT and the technical support group personnel. 

In closing this section of the paper there are some general observations that should made 

regarding the multiple benefits of serving as a TDT member. Working as a member of an NESC 

TDT clearly offers challenging opportunities. Members of the TDT interact with the best of the 

best in NASA, industry, academia, and other government agencies to address a broad spectrum 

of discipline technical issues. TDT members also find that working within the NESC 

organizational structure permits an exposure to other NASA programs, projects, cultures, 

methods, and business practices from across the Agency. Typically, this allows experiences to be 

gained outside one’s normal work area within a single NASA Center organization. The 

experience should broaden one’s horizons via the wide network of job-related interactions. There 

will be technically challenging and diverse assignments of a high impact/high feedback/high 

visibility nature. Serving on a TDT provides an avenue for both professional growth and positive 

recognition, not only within the discipline Community of Practice but also within the NESC’s 

customer and stakeholder community. The overall TDT experience is one that should provide 

motivated, tenacious, and intellectually curious team members with a very high degree of job 

satisfaction. 

4.0 Experiences of the NESC GN&C TDT 

The GN&C TDT has engaged in multiple NESC assessments, consultations and reviews since 

the NESC became an operational organization in November of 2003. Reference 1 provides a 

description of NESC GN&C TDT activities over the period 2003 to 2007. 

Highlighted in this section are several of the most recent TDT experiences. In certain cases, 

significant relevant technical details are provided as background. These experiences, which 

occurred over the period from 2007 to present, were selected to illustrate the wide variety of 

work the GN&C TDT engages in. The reader will see that the GN&C TDT supported both 

human space flight projects and robotic spacecraft projects. The time durations over which the 

work was performed varied from durations of a few weeks to a month or two for the smaller 
6 

scale quick-reaction peer review tasks to durations of several months to over a year for larger 

scale efforts. In cases where the task was primarily focused on a specific GN&C discipline issue, 

the work was performed exclusively by the NASA Technical Fellow for GN&C and/or small 

contingents of GN&C TDT members. In other cases that required a more multi-disciplinary 

approach the GN&C TDT members supported the task as part of a larger, integrated NESC team 

effort under the direction and leadership of an NESC Project Engineer. As with the majority of 

NESC endeavors, experts from virtually all NASA Centers, other government Agencies, national 

laboratories, academia, and industry were involved in conducting these activities. 

4.1 Max Launch Abort System (MLAS) Flight Test Assessment 

The NESC designed, developed, and flew the alternative MLAS as risk mitigation for the 

baseline Orion spacecraft launch abort system (LAS) already in development. The NESC was 

tasked with both formulating a conceptual objective system (OS) design of this alternative 

MLAS as well as demonstrating this concept with a simulated pad abort (PA) flight test. The 

goal was to obtain sufficient flight test data to assess performance, validate models/tools, and to 

reduce the design and development risks for a MLAS OS. 

In June 2007, the Associate Administrator for the Exploration Systems Mission Directorate 

(ESMD) requested the NESC undertake the MLAS Project. The MLAS was named after 

Maxime (Max) Faget. Dr. Faget was the lead designer of the Mercury space capsule and 

developed its abort system called the “Aerial Capsule Emergency Separation Device”. It was in 

his honor that the MLAS was named. His innovative spirit and his team’s rapid development of 

new technologies formed the inspiration for the MLAS Project. 

The charter for the MLAS Project was to develop, design, and test an alternate concept for the 

Orion Crew Exploration Vehicle (CEV) LAS. MLAS would be theoretically capable of 

extracting the Orion vehicle from the launch vehicle at any time from crew ingress at the launch 

pad through staging and ignition of the second or upper stage of the Ares-I crew launch vehicle. 

The MLAS Project would conclude with at least one full-scale unmanned PA test suitable for 

demonstrating the viability of this alternate LAS concept. The MLAS project would be run 

independently from the Constellation Program (CxP) and Orion Project in order to minimize 

impact on in-line program resources. It was dictated that off-the-shelf hardware and existing 

technology would be used wherever possible on MLAS. Design and development work 

previously accomplished by the NESC would be leveraged for this project. Previous work 

includes the CEV smart buyer design (SBD), composite crew module, and the alternate launch 

abort system. This NESC infrastructure was used to form the NASA-wide MLAS Project team. 

For example, members of the GN&C TDT (Reference 1) were recruited to serve on the MLAS 

GN&C team and other members were subsequently recruited to serve as peer reviewers of the 

GN&C team’s work. 

The origins of the tower-less MLAS flight test vehicle (FTV) can be traced to the initial notional 

‘back of the napkin’ drawing (see Figure 1a) that was conceived during a CEV SBD study 

outbrief brainstorming session in March 2006. Subsequently, a refined MLAS drawing (see 

Figure 1b) was provided to NESC in June 2007 by Scott “Doc” Horowitz, the Associate 

Administrator for the ESMD at the time MLAS was initiated, as a notional point of departure for 

the early system level trades and vehicle configuration studies. It is interesting to compare the 

final MLAS FTV prelaunch physical configuration (see Figure 1c) with the original notional 
7 

drawings depicted in Figure 1a and Figure 1b. One can clearly see that the MLAS concept differs 

from Orion LAS in that it uses side-mounted abort motors instead of a tower-tractor abort motor 

design that pulls the crew module (CM) from above. The tower abort motor has been designed 

out of the MLAS concept and the CM fully encapsulated in a forward fairing. 

a) First notional 

conceptual drawing 

(2006). 

b) Refined conceptual drawing 

(2007). 

c) Mounted on launch stool at 

Wallops Island (2009). 

Figure 1. MLAS FTV. 

The MLAS GN&C team was formed in July 2007 and was given the responsibility for modeling, 

simulating and analyzing the trajectory and attitude dynamics of the MLAS FTV during its 

simulated PA flight. The GN&C team ensured that the flight test occurred within the envelope 

defined by the requirements, and they constructed the nominal target flight timeline and 

trajectory. The GN&C team selected all of the MLAS trajectory and attitude flight 

instrumentation equipment including the inertial measurement units (IMUs) and global 

positioning system (GPS) receivers. The team chose these navigation sensors to generate flight 

test data that would permit post-flight reconstruction of MLAS vehicle trajectory and attitude 

dynamics. Analyses performed by the GN&C team determined ballasting, motor alignment, and 

launch stand angle requirements. Over the course of the MLAS Project the GN&C team worked 

closely with the System Engineering and Integration, Aerodynamics, Landing and Recovery 

System, Propulsion, Avionics, Software, Structures, and Loads and Dynamics teams and also 

with the MLAS Chief Engineer. The MLAS Project was by far the most extensive activity 

undertaken by the GN&C TDT over the period of time being covered in this paper. 

The MLAS simulated PA flight test was successfully conducted from Wallops Island on 8 July 

2009. Figure 2 shows the MLAS FTV shortly after abort motor ignition. Figures 3 and 4 depict 

the CM altitude versus range and the CM altitude versus time, respectively. Figures 3 and 4 show 

the position data from the GPS receivers alongside the pre-flight predictions made with the wind, 

density, static pressure and speed of sound data from the L-3 minute balloon sonde. The as-flown 

flight trajectory was somewhat steeper than that predicted in the pre-flight simulation; however, 

the trajectory deviation from the prediction was within 0.75 standard deviations in all directions. 
8 

Figure 2. Launch of MLAS FTV from Wallops Flight Facility (8 July 2009). 

Figure 3. MLAS CM altitude versus range. 
9 

Figure 4. CM altitude versus time. 

The GN&C TDT can look back upon the MLAS experience and observe that less than two years 

after project start the flight test of this alternative launch vehicle abort system was conducted in 

very close agreement with its operations concept. All the key flight test events occurred in the 

prescribed sequential order and within acceptable tolerances of the exact pre-planned time and 

flight dynamic conditions. The entire flight test duration was 88 seconds during which multiple 

staging events were performed, and nine separate critically timed parachute deployments 

occurred as scheduled. Overall, the as-flown flight performance was as predicted prior to launch. 

The primary lesson learned on MLAS was the extent to which the design and integration of a 

LAS can impact and influence the overall launch vehicle/spacecraft system design, overall 

reliability, and performance capabilities. Several other important lessons learned from this 

MLAS flight test experience are: 

 The need for all project team members to look beyond their immediate discipline task to 

think, speak up, and act like Systems Engineers for the benefit of the project. 

 The need for planned periodic crosschecks of critical analytical results by having 

technical ‘shoot outs’ between different engineering groups using different tools/methods 

to independently generate analytical products. 

 The need for periodic and informal face-to-face peer reviews; the emphasis should be on 

reviewing the details of modeling assumptions, analytic methods, uncertainty 

assumptions, simulation results, control law algorithm designs, software code, etc., and 

not on preparing formal presentations and responding to action items. 
10 

 The need for stringent mass properties testing, and vehicle mass properties configuration 

control measures and routine periodic reports, especially on a passively controlled 

vehicle. 

 The degree of difficulty in safely and reliably performing precision alignment of the solid 

rocket abort motors was initially underestimated. 

 The need to manage and limit insidious unchecked vehicle mass growth which erodes 

flight performance, degrades system design robustness, increases parachute loads and 

generally diminishes the probability of overall mission success. Associated with this is 

the need to incorporate realistic growth margins into the vehicle’s design to mass budget. 

In addition, this can be mitigated in part by defining a robust set of structural loads 

requirements as early as possible in the project design cycle. 

 A rapid prototyping activity like MLAS can be accomplished using virtual meeting 

technologies supplemented by periodic Co-Locations of the entire team and by 

leveraging modern online concurrent data sharing and configuration management 

capabilities. 

4.2 Constellation System Preliminary Design Review (PDR) Assessment 

The Constellation Chief Engineer requested the NESC to participate in the CxP Preliminary 

Design Review (PDR) process as the Independent Review Team (IRT). The NASA Technical 

Fellow for GN&C, along with members of his GN&C TDT, served on this IRT. The purpose of 

the IRT was to ensure technical issues were not overlooked due to familiarity with the PDR 

products. The IRT was asked to identify technical issues in three categories. The first category 

was an issue or resolution to an issue that was incomplete or not fully addressed. The second 

category was an issue that was not addressed. The third category was an issue that the IRT 

disagreed with its resolution or had a contrary opinion. 

The CxP PDR process, which culminated in April 2010, had the following several objectives: 

 Demonstrate the End-to-End (E2E) understanding of the Initial Operational Capability 

Mission and where the design is driven by the Lunar Mission 

 Demonstrate that the completed analysis and tests provide the results necessary to close 

the architecture to accomplish the Design Reference Missions 

 Demonstrate the architecture performance is capable of accomplishing the E2E mission 

with acceptable risk 

 Identify key architecture risks and associated mitigation status 

 Baseline Initial Capability Interface Requirements Documents and draft Spacecraft 

Interface Control Documents 

 Documented Program Test & Verification Strategy and draft System Integration Plans 

released 

 Define the Forward Work plan necessary for a successful CxP Critical Design Review 
11 

The IRT members first participated in a series of discipline-specific Face-to-Face Integration 

Reviews, each led by the respective phase lead engineer. This was an excellent approach to 

somewhat informally ‘break the ice’ technically in each of the mission phases (e.g., ascent, onorbit, entry descent and landing, etc.) prior to the actual formal CxP PDR milestone. These 

mission phase reviews were held near the Johnson Space Center (JSC), Houston, Texas, over the 

period from November 2009 through December 2009. In addition, the IRT participated in the 

PDR kickoff meeting held the week of January 12, 2010 and the PDR Pre-Board and PDR Board 

meetings held the week of March 1, 2010. The IRT provided 41 review item discrepancies (RID) 

and five comments in the RID system. 

Overall, the GN&C Technical Fellow and the participating members of the GN&C TDT believed 

the CxP PDR process was a good high-level review of the Program maturity and identified the 

issues, work completed, work to be performed, etc. The face-to-face meetings created an 

environment of open communication among all participants. The face-to-face teams were very 

receptive to the IRT’s comments and met separately with the IRT to alleviate concerns and 

issues. The RID Review Teams (RRT) attempted to understand the issues by speaking directly 

with the RID initiators rather than just reviewing the RIDs on paper. The RRTs tried to 

accommodate the initiators’ schedules and coordinated with them to be able to state their case 

directly. When impasses were reached, the reclama process was recommended to the initiators. 

The IRT agreed that the RRTs had worked with all the RID initiators to reach mutually agreeable 

resolutions to the issues that were raised and all RIDs that the IRT submitted were acceptably 

resolved. 

Clearly, the PDR process was a very successful forcing function to bring forth “gaps” and 

discrepancies. The identification of these occurred because of the open communications with the 

IRT and between all levels of the program and projects. However, it was apparent that not all the 

needed integration was occurring during normal program functions prior to the PDR process. It 

is understandable to have elements that lag behind others when everyone is concentrating on 

their top issue or latest problem. New issues were raised and clarifications were generated during 

the PDR process. For example, a RID was written to address the fact that, as presented, the AresI Roll Control System (RoCS) propellant margins (at about 8%) were lower than the 20% margin 

typically desired at the time of the system PDR design milestone. 

The issues that were identified by the IRT were grouped into four broad themes and summarized 

in a NESC White Paper provided to CxP program managers and others. The items discussed in 

the NESC White Paper fell into all three categories of technical issues that the IRT was charged 

to identify. The first theme identified issues that were lacking the technical maturity required for 

a PDR. The areas of launch abort vehicle (LAV) performance in the transonic flight regime, 

updating of the integrated Probabilistic Risk Analysis (PRA) model, regression test planning, 

incorporation of human factors to support design trades and studies, and the mismatch of 

program and project maturity levels fell within this category. The recommendations included 

pursuing engineering solutions at this early stage of the design process and saving operational 

changes for later as the program matures and when design changes can be prohibitively 

expensive to implement. 

The second theme addressed uncertainty in models and simulations. The GN&C Technical 

Fellow and the GN&C TDT members felt the interpretation, quantification, and application of 
12 

model and data uncertainty were especially important areas to be addressed by both CxP 

managers and engineers. It is acceptable to have uncertainty in models and simulations, but 

quantifying the uncertainty and clearly communicating the model uncertainty to perform trades 

and to make informed decisions of risk needs to be emphasized. 

Mathematical models, and their associated data sets, are used to capture and represent the current 

state of knowledge of system physical behaviors in every element of the CxP architecture. These 

models and their supporting data are exercised to analyze and predict system sensitivities, 

performance against requirements, factor of safety/stability margins, risks and overall mission 

success. 

It is a recognized and accepted reality that these models and their data have limitations. Some 

models have insufficient experimental data to characterize precisely a probability distribution of 

critical model parameters and features. In other cases, there is only poor or limited understanding 

of the underlying physics phenomena or physical coupling between system elements or between 

sub-elements. Beyond our inability to completely understand the dynamics of a given problem 

there is also the reality that there will be both variability in the physical features/parameters of 

the as-built flight systems and additional variability in the operational environments encountered 

during manufacturing and assembly, ground operations, and mission operations. All of these 

aspects need to be accommodated and quantified in an uncertainty framework by the engineers 

using the models/data. 

Various methods and approaches have been used across CxP and its elements to characterize the 

uncertainty in these models and their supporting data. The NASA-STD-7009 for Models and 

Simulations was developed to help NASA programs/projects consistently address the challenges 

of documenting, analyzing, and communicating uncertainties, and several other modeling and 

simulation issues in the Agency. The NASA-STD-7009 also requires clear documentation and 

communication of model assumptions to decision makers. Reporting uncertainties in model 

inputs and outputs is required for critical analyses, and data/model uncertainty quantification are 

factors in the credibility assessment scale (CAS) review. The CAS review provides a common 

framework to evaluate models and simulations. 

However, there is not a consistent interpretation of uncertainty across the CxP architectural 

elements, and NASA-STD-7009 does not address inconsistency in model uncertainties across a 

set of analyses within a program. This is a concern since having a proper and consistent 

application of model/data uncertainties is fundamental to providing CxP decision makers with 

informed risk analysis information. Consistency is especially important where CxP decision 

makers are seeking to trade risk across architectural elements, across system elements, and 

across technical disciplines. Furthermore, flight system designs can be excessively overconservative if modeled uncertainties are significantly higher than the variations or unknowns of 

the physical system. Conversely, designs can be at risk of failure if modeled uncertainties are 

lower than the actual variations of the physical system. Input data uncertainties are included in 

the integrated loss of crew/loss of mission models, and the PRA results include reporting 

uncertainty ranges using 5th and 95th percentiles. 
13 

An example of the model uncertainty issue is a recent aerodynamic database model uncertainty 

distribution change that resulted in a reduction of the LAS attitude control motor (ACM) control 

authority in the transonic flight regime (M = 0.9–1.3, 35–47 sec mission elapsed time). This has 

driven the need for the CxP to investigate alternative means of improving performance of the 

LAV (e.g., adding ballast mass). In parallel, special wind tunnel tests are being pursued in an 

effort to lower the uncertainties in the aerodynamic database. 

There is some level of recognition that the application of model/data uncertainties needs to be 

improved within CxP and certain teams are actively working the issue. It is not clear however 

that the current practice of inconsistent interpretation and application of uncertainty across CxP 

has been elevated sufficiently high and captured as an integrated risk for the program. A 

common uncertainty framework for all CxP elements that is consistent with NASA-STD-7009 

would provide needed consistency and rigor in addressing the issues associated with model/data 

uncertainty. Per NASA-STD-7009, reporting uncertainties in model inputs and outputs is 

required for critical analyses and data/model uncertainty quantification, documentation, and 

communication are factors in the CAS review. 

With all the great work in models and simulation, a trap can occur where there is an overreliance 

on models and simulation. This can occur because the people who are doing the day-to-day 

problem solving are so caught up in working with the models and trying to reduce and define the 

uncertainty associated in the models that they fail to realize the solution they are obtaining can 

defy the basic laws of physics and engineering. Using engineering judgment to check the 

analytical results, engaging the critical end users in the process of quantifying uncertainties, and 

periodically re-evaluating assumptions and processes should occur to avoid this trap. 

The third theme addressed in the NESC White Paper deals with disconnects in the level of 

maturity between project elements and interfaces. The last theme discussed the process of 

identifying test needs. The IRT noticed a perception that some testing was not being pursued due 

to reasons of cost or schedule even though the test might significantly buy-down risk. The IRT 

recommended the development of a list of minimum required tests (i.e., an ‘incompressible’ test 

list). 

This particular experience reminded all the teams involved that an independent review can 

provide a fresh perspective not biased by long-term involvement with the project or program. 

The Face to Face Integration Reviews, led by the respective phase lead engineers, was a 

tremendously valuable first step in the overall CxP PDR process. For the CxP PDR, the IRT 

worked with the program and project personnel to identify new issues or issues that the program 

knew of, but did not have the right management visibility. More importantly, the IRT provided 

recommendations that should be pursued. 

4.3 Orion Ascent Abort One (AA-1) Loads Assessment 

In April 2009, the Systems Engineering Integration Team Lead in the Orion Test Flight Office 

requested that the NESC perform an independent assessment of AA-1 coupled loads analysis 

performed in support of AA-1 Load Cycle-1(LC-1). 
14 

Based on data from AA-1 LC-1, there were concerns regarding the high-load levels predicted 

during vehicle separation, twang, ascent, and maximum dynamic pressure (Max-Q). The NESC 

Team was tasked to perform an independent review of data and provide consultation on the 

analysis approach and associated assumptions, limitations, and uncertainties. The intent of the 

NESC review was to thoroughly scrub all aspects of LC-1 and then provide findings, 

observations and recommendations for any changes to be incorporated into the follow-on effort, 

LC-2. The NESC Team was not expected to provide an independent load analysis unless 

significant issues were identified. 

Figure 5 provides an illustrative comparison between the PA-1 flight test configuration and the 

AA-1 flight test configuration. The AA-1 LAV system includes the LAS, LAS fairing assembly, 

and CM. The AA-1 abort flight test will launch from the White Sands Missile Range. During the 

boost phase, the LAV will be connected to the abort test booster (ATB) through the separation 

ring (SR). The ATB will boost the LAV to an altitude of nearly 38,000 feet before separation of 

LAV from ATB/SR. The LAV abort motor (AM) and ACM will ignite and burn for less than 

five seconds. Then the LAS will separate from the CM, and the CM will descend via parachute 

to a soil landing. 

The critical loads for the LAV occur during the motor burn, parachute deployment, and landing. 

However, LC-1 covered only the motor burn phase, which included LAV separation from the 

ATB, ignition of the AM, and flight up to Max-Q. 

Several members of the GN&C TDT, together with members of the NESC Flight Mechanics 

TDT, supported this assessment by investigating and reporting on the technical features of the 

ATB boost phase flight control system (FCS) and other aspects of the AA-1 GN&C system 

design that influenced the vehicle loading during flight. In particular, the GN&C and Flight 

Mechanics TDT members concentrated on how the Monte Carlo analysis of the AA-1 boost 

phase and resulting separation conditions was performed. 

The team determined that LC-1 was adequate as a preliminary set of initial design loads for AA1 considering the number of uncertain or undefined parameters in structure, aerodynamics, event 

timing, etc. that existed at the time of the analysis. Given these qualifiers, it is not surprising that 

numerous areas were noted where improvements could be effected in LC-2. 
15 

Figure 5. Comparison of PA-1 and AA-1 flight test configurations. 

Some of the items the NESC team found necessary to be corrected or implemented for LC-2 to 

produce a complete set of detailed design loads and thorough structural assessments included: 

 Develop a Project Technical Requirements Document that specifically defines AA-1 

hardware functional and design requirements, environmental requirements, interfaces 

with other hardware, etc. 

 Document the use of uncertainty factors or dispersions in the various LC-2 inputs to the 

analysis for concurrence by the responsible technical panels and the Orion Flight Test 

Office.

 The thrust alignment/offset dispersions of the abort motor should be clearly defined (1 

degree was arbitrarily used). 

 The AM thrust offset dispersions to be used in the in LC-2 analysis should include such 

factors as temperature distortions, assembly tolerances, as alignments. 

 The boost phase Monte Carlo analysis should be used to define the required range of 

initial flight conditions at separation for the loads analysis in LC-2. 

 The full specified range of allowable center of gravity (CG) locations should be explored 

in LC-2. 
16 

 A sensitivity analysis should be performed to check the effect of extreme “corner of the 

box” CG mass properties on modal characteristics. 

 Propellant mass changes should be included in the LC-2 analysis as those changes affect 

both structural dynamics and CG shifts. 

4.4 Orion CM Boilerplate Water Drop Test Assessment Support 

In October 2009, the NESC initiated an assessment to perform the specific tests and analyses 

needed to provide the necessary verification to correlate LS-DYNA® simulation results with data 

collected during multiple water drop tests of a full-scale Orion CM boilerplate test article. The 

LS-DYNA® code is an advanced general-purpose multiphysics simulation software package 

developed by the Livermore Software Technology Corporation and used to design the Orion CM 

structure. The LS-DYNA® software package, which employs finite element analysis methods 

using explicit time integration, is especially well suited for the simulation of highly nonlinear 

transient dynamic events such as the impact dynamics of the CM structure when it lands in 

water. The nonlinear effects typically manifest themselves as changing boundary conditions, 

relatively large material deformations, and certain materials that do not exhibit ideally elastic 

behavior. Transient dynamic means analyzing high-speed, short-duration impact events where 

inertial forces are important. 

The fundamental structural design simulation tool for the Orion Project was LS-DYNA®, 

specifically developed to accommodate non-linear load paths as occur during water impacts. The 

NESC Assessment Team surveyed readily available LS-DYNA® literature on the use of the tool 

to model water impacts, and assessed that some of the techniques used are not deterministic but 

rely on manipulation by the analyst. Experimental data was required to anchor LS-DYNA®

analyses of water impacts to mitigate the risk of the sort of impact damage seen during Apollo 

drop tests. A strong motivation for these NESC-sponsored tests is the very limited number of 

landing impact tests in the Orion Project plan planned to support the CM primary structure 

development and qualification. The Orion Project had concerns regarding the possibility that 

significant uncertainties would remain in the landing load environment, resulting in an unknown 

structural capability and an inability to achieve structural certification for flight. Since LSDYNA® is being used to establish the structural design of Orion, this program and other NASA 

programs are critically dependent on the validity of the answers generated by the tool. This 

independently performed and fast-paced series of drop tests, and the subsequent detailed 

evaluation of the data collected, are intended to support risk determination for the Orion Project 

and NASA, potentially influencing future CM testing plans and design options. This particular 

NESC assessment is contributing to the ultimate NASA goal of developing an integrated landing 

system design, with a mass-optimized space capsule structure, that will attenuate the impact 

loads transmitted to the crew to non-injurious levels. The key stakeholders for this assessment 

are the Orion Project team and the NASA Office of Chief Engineer. 

The NESC’s MLAS CM simulator (boilerplate) was identified as a rapidly available test article. 

This test article was instrumented in two phases, and tested via crane drops into water. 

Instrumentation included multiple inertial sensors, photogrammetric targets, pressure 

transducers, and strain gages. The first phase of drop tests were performed at the U.S. Army’s 

Aberdeen Proving Ground located in Maryland. The test article was hoisted via a crane to a predetermined level and then released to drop into the water (see Figure 6). In Phase I of this drop 
17 

test assessment, a total of eighteen (18) drops were conducted: fourteen (14) were dropped from 

a height of ten feet and the remaining four (4) were “low-energy” drops released at a height of 

eight (8) feet. 

Members of the NESC GN&C TDT were requested to support both the generation of quick-look, 

in near realtime, inertial sensor data products as well as to perform the more detailed post-test 

processing and analysis of the water drop test data generated by multiple inertial sensors. The 

GN&C TDT members augmented the existing drop test by bringing the necessary skills and 

expertise to understand, convert, transform, format, process, and analyze data from the multiple 

gyroscopes and accelerometers mounted on the test article. The inertial sensors included four triaxial accelerometer packages, two packages mounted very near the CM test article’s CG and the 

other two mounted near the apex of the test article’s apex. There were also individually mounted 

analog devices single-axis micro electro-mechanical systems angular rate sensors (gyroscopes) 

as well as a Humphreys Crossbow® IMU. These inertial sensors were mounted on an aluminum 

structural crossbeam (see Figure 7). 

Figure 6. Orion CM boilerplate water drop test at Aberdeen Proving Ground. 
18 

Figure 7. Inertial sensor instrumentation mounted inside CM test article. 

Acceleration data was collected using analog sensors with a bandwidth of 1000 Hz, which was 

then sampled at a rate of 13,888 Hz, embedded in a telemetry stream, and recorded. A 

representative example of the raw data collected is the longitudinal axis (X-axis) acceleration vs. 

time data plot is presented in Figure 8. Quick-look analysis of the acceleration data collected 

revealed a noise floor at approximately –15 dB along with the undesired presence of several 

unwanted CM structural modes, in particular the structural modes of the aluminum crossbeam on 

which the sensors were mounted. The flexible body modes can be seen in Figure 8 over the range 

from 100 Hz – 1000 Hz with the peak flexible mode response occurring at 375 Hz at 

approximately +8 dB. A proper analysis of the acceleration data therefore required the 

attenuation of unwanted high-frequency structural “noise” while keeping the low-frequency test 

article rigid body information and the all-important relatively short duration (150 milli-second) 

splashdown transient. So, in addition to performing all the necessary data conversions, reformatting, and coordinate transformations of the various inertial data sets collected, the GN&C 

TDT contingent spent a great deal of time and effort studying and eventually selecting the type 

and order of digital filter that would sufficiently attenuate the high-frequency structural ringing 

effects from the desired rigid body motions. 

A third-order low pass filter (LPF) was initially selected to filter the raw accelerometer data 

collected. This LPF had two second-order poles each at 75 Hz and a single first-order pole at 85 

Hz. The second-order poles had a 0.7 damping ratio. The continuous form of the LPF (see 
19 

Equation 1 for this representation) was then converted to a digital LPF via the standard bi-linear 

transformation using a sampling period of 73.5 micro-seconds. 

 (1) 

Figure 9 shows the result of digitally low pass filtering the raw X-axis accelerometer data. 

Clearly, the flexible body effects are significantly reduced leaving a signal more closely 

representing the CM rigid body response. After much further technical deliberation with the LYDYNA® users a 6th-order LPF, applied in one direction only, was selected by the GN&C team to 

perform the necessary attenuation. Another significant effort concerned determination of the test 

article’s pre-drop initial conditions. Photogrammetric data was utilized to both define the initial 

conditions and the gyroscope drift and bias estimates. The GN&C TDT members provided 

sufficient processed inertial sensor test data to permit the comparison and verification of the LYDYNA® simulation results. Phase II water drop test with improved inertial instrumentation are 

planned for early FY2011. 

 

Figure 8. Representative raw X-axis CG#1 accelerometer data from CM water drop test. 
20 

Figure 9. Low pass filtered X-axis CG#1 accelerometer data from CM water drop test. 

4.5 Assessment of Orion Slosh Impact on Ares-I FCS Stability 

The deleterious effects of propellant sloshing on launch vehicle stability have been long 

recognized by FCS designers. Slosh dynamics were carefully considered on both the Apollo 

Saturn-V launch vehicle and the Space Shuttle launch vehicle. The term “sloshing” is used here 

to describe the phenomenon of the free surface oscillations of the fluid (either fuel or oxidizer) in 

a partially filled tank. 

In December 2009 the NESC GN&C TDT was asked to assess the impacts of the Orion Service 

Module (SM) slosh dynamics on the Ares-I crew launch vehicle FCS stability and performance. 

Figure 10 depicts the Ares-I launch vehicle shortly after liftoff from Kennedy Space Center. The 

NESC assessment team sought resolutions to two fundamental issues: 1) assess conservatism in 

Marshall Space Flight Center’s (MSFC) Ares-I GN&C impact analysis of Orion SM bare tank 

slosh dynamics; and 2) Determine if the MSFC GN&C team can design an Ares-I FCS, which 

accommodates the 0.03% Orion SM bare tank slosh mode damping and satisfies all stability 

robustness and performance requirements. Given the short time duration of this assessment, the 

NESC did not perform any independent GN&C modeling, simulation, or analysis. Rather the 

focus was placed on understanding the details of the Ares-I GN&C modeling, simulation and 

analyses performed by the MSFC flight control engineers. 
21 

Stabilizing the Orion SM slosh mode dynamics on top of the Ares-I flexible-body and LOX/LH2

slosh dynamics poses a challenging flight-control design problem for MSFC. Extremely low 

values of Orion slosh damping emerged as early as June 2009 as a key driver in the Ares-I 

GN&C design. The NESC first became fully aware of the sensitivities of the Ares-I FCS design 

to the extremely lightly damped Orion slosh dynamics in November 2009. This occurred at the 

CxP Ascent/Ascent Aborts Integrated Phase Review held at JSC as a precursor to the CxP 

System-level PDR board. This was true cross-project system engineering and integration issue 

between Orion and Ares. At that time, as a minimum, it was recommended that a new CxP 

Program-level risk be created and tracked within the CxP Integrated Risk Management 

Assessment (IRMA) framework to provide visibility of the Orion/Ares slosh problem. Several 

elements of risk were to be included in this new IRMA risk item, including: 

 Risk to Ares GN&C resources and schedule, 

 Lack of flexibility to accommodate future Ares vehicle challenges, changes, and 

opportunities, 

 Loss of Ares capability to accommodate failures, 

 Risk for meeting program stability and margin requirements, 

 Risk for increased structural loads, 

 Loss of significant robustness and hence safety, and 

 Risk to Ares’ ability to certify GN&C for flight readiness. 

Figure 10. Ares-I launch vehicle shortly after liftoff. 
22 

The challenge for the Ares-I FCS designer is to meet simultaneously the stability margin 

requirements cited above and the linear step response requirements. There are other high-level 

flight performance requirements that also must be satisfied but these were the one of most 

interest to the NESC assessment team. Lastly, it should be noted that no minimum fluid slosh 

damping requirements were ever placed, by the Ares-I Project, on the Orion Project’s SM fuel 

and oxidizers fluid/tank system. 

The Ares GN&C 7 design cycle analysis, completed by MSFC flight control engineers in June 

2009, had indicated Ares-I FCS sensitivities to Orion SM fuel and oxidizer slosh dynamics. The 

problem was attributed to two major factors: 1) the Orion’s SM fuel and oxidizer tanks are 

partially filled, resulting in a significant sloshing mass; and 2) the Orion SM tanks have very 

smooth bare walls. Slosh mode damping values, determined using the industry-standard NASA 

SP-106 handbook (Reference 2) methodology were computed to be an order of magnitude below 

NASA heritage. The NESC team performed a comparison of the Orion SM slosh damping values 

with the Ares-I launch vehicle slosh damping values and legacy slosh damping values from the 

Space Shuttle Program and the Apollo Saturn launch vehicle. The empirically computed 1-g 

minimum slosh mode damping value for the Orion nitrogen tetroxide (NTO) oxidizer tank was 

0.0003 or 0.03%. Likewise, the empirically computed 1-g minimum slosh mode damping value 

for the Orion monomethylhydrazine (MMH) fuel tank was 0.0006 or 0.06%. The NESC 

assessment team verified these empirical 1-g calculations of damping values. 

The MSFC Ares-I flight control team was unable to develop a design that would meet stability 

robustness requirements using the 0.03% (NTO) and 0.06% (MMH) damping values. The Ares-I 

flight control team was given action to quantify impacts to their FCS design due to the extremely 

low Orion SM tank damping values. Subsequently multiple Ares-I FCS redesigns were 

performed to accommodate partially filled, low-damped Orion SM tanks. In order to achieve the 

required stability margins for both first stage and upper stage flight in the GN&C 7 Ares-I design 

cycle the Orion SM tank slosh damping values were arbitrarily increased from the empirically 

calculated smooth wall 0.03 – 0.06% damping values to a damping value of 0.5%. At the time 

the NESC assessment was initiated these Ares-I analyses were only in a preliminary state and it 

was recognized that another full Ares-I GN&C design cycle (i.e., the GN&C 8 cycle) would be 

required to fully quantify the flight control issues and impacts. In particular it was noted that the 

impacts on the slosh dynamics due to many routine transient events such as liftoff “twang”, stage 

separation, LAS jettison, SM panel deployment, the ullage settling motor firings, and potential 

interactions with the non-linear reaction control system roll axis controller, had not yet been 

studied. 

As depicted in Figure 11, the FCS architecture for the Ares-I consists of a classical proportionalintegral-derivative (PID) controller design with loop shaping compensation (i.e., digital filters) to 

provide phase and gain stabilization of the flexible body modes and/or fluid slosh modes. The 

“proportional” term determines the controller response to the currently sensed error, the 

“integral” term determines the response based on the sum of recent errors, and the “derivative” 

term determines the reaction based on the rate at which the error has been changing. The 

weighted sum of these three different control actions is used to generate the thrust vector control 

(TVC) command for the next control cycle. The PID controller gains and the filter parameters 

are all gain scheduled which is consistent with industry practice. 
23 

The pitch and yaw axes TVC system controllers share this common system architecture. The 

Ares-I roll axis controller takes a different form and is a non-linear phase-plane RoCS thrusterbased controller quite different from the TVC pitch and yaw controllers. The roll controller has 

been developed from legacy Shuttle and International Space Station ISS phase-plane controller 

designs. 

Table 1 summarizes the Ares-I FCS stability gain and phase margin requirements for the lowfrequency rigid body/aero/slosh condition as well as for the cases where the flexible body modes 

of the launch vehicle are either gain or phase stabilized. Traditionally, the gain and phase 

margins are used as control system stability robustness metrics and they represent a tolerance to 

both plant disturbances and environment uncertainties. The requirements shown in Table 1 are 

consistent with industry practice. Note that under fully dispersed conditions the minimum rigid 

body/aero/slosh gain and phase margin requirements are suitably lowered to 3 dB and 20 

degrees, respectfully, but are not dropped to zero as done on some projects. The establishment of 

diminished but finite dispersed gain and phase margins in this manner is especially prudent given 

the developmental nature of the Ares-I launch vehicle. These Ares-I stability margin 

requirements listed in Table 1 are very similar to, and one can most likely trace their origin to, 

the stability gain and phase margin requirements levied on the Space Shuttle launch vehicle 

which are shown in Table 2. 

Figure 11. High-level Ares-I FCS architecture. 
24 

Table 1. Ares-I FCS stability margin requirements. 

Table 2. Space Shuttle launch vehicle FCS stability margin requirements. 

Parameter 

Condition 

Nominal Dispersed 

Rigid Body GM (dB) 6 3 

Rigid Body PM (deg) 30 20 

Slosh Mode PM (deg) 25 15 

Slosh Mode GM (dB) 6 3 

Flexible Body Mode Attenuation (dB) 10 6 

The Ares-I gain and phase margin definitions for rigid body, slosh, and flex body dynamics are 

notionally illustrated in Figure 12 (first-stage example) and Figure 13 (upper-stage example). 

The red regions reflect ±6 dB gain and ±30-degree phase from the critical stability point. Lower 

frequency gain margins (rigid/aerodynamics and slosh) are defined at the 180-degree phase 

crossover, while required high-frequency flex gain attenuation is defined as peak gain, measured 

from the 0 dB axis. For Ares-I phase stabilized flexible modes, both lead and lag phase margins 

are defined. The general design philosophy for Ares-I first-stage flight is to robustly phase 

stabilize the first flexible mode, and robustly gain stabilize the second and all higher flexible 

modes. The design approach for upper-stage flight control is to have all vehicle flexible modes 

be gain stabilized. 
25 

Figure 12. Notional depiction of Ares-I FCS gain and phase margins (first-stage example). 

Figure 13. Notional depiction of Ares-I FCS gain and phase margins (upper-stage example). 
26 

Figure 14 depicts, for a specific ISS mission case, the control system, flexible body, and slosh 

mode frequencies as a function of time, with uncertainty included, over the period from launch to 

600 seconds after launch. This figure graphically illustrates the challenges of designing a control 

system for first-stage powered flight. The launch vehicle flexible body dynamics must be 

stabilized by the FCS throughout flight. A large percentage of the vehicle mass is composed of 

propellant, which is consumed during the boost phase. This loss of mass significantly changes 

the structural dynamics as propellant is depleted. For most fill levels, slosh frequency scales 

proportionally with vehicle acceleration and damping scales inversely. The Ares-I first-stage 

maximum acceleration is about 3 g’s, which results in increased slosh frequencies and decreased 

damping. Stability issues can occur with this shift in Orion slosh frequency under the influence 

of the launch vehicle 3-g acceleration. 

The potential for destabilizing Ares-I flexible body mode dynamic coupling and interaction with 

Orion NTO and MMH slosh modes is apparent from an examination of Figure 14. The Ares-I 

full-stack structure first flexible mode frequency is very near the same frequency as the Orion 

SM slosh modes, which can cause decreased margins and difficulty in filter design. The MSFC 

team pointed out that the NASA SP-8031 “Slosh Suppression” guidelines document (Reference 

3) recommends separation between the flexible body and fluid slosh mode frequencies to 

minimize flex body/slosh dynamic interaction. 

During this assessment, questions were raised by the GN&C TDT members regarding the 

implementation details of MSFC’s linear spring-mass slosh mechanical model used in their 

Marshall Aerospace VEhicle Representation In “C” (MAVERIC) 6-degree of freedom (DOF) 

simulation. There was discussion comparing linear versus non-linear slosh modeling. This topic 

arose after the NESC team requested to see output data on the slosh amplitudes seen in MSFC 

simulations. The need for and the potential benefits of a non-linear slosh model were discussed

Plans for an Ares-I GN&C mini-design cycle where the stability margin and performance 

sensitivities will be evaluated over a range of Orion damping values from 0.1% to 0.5% were 

also discussed. The stated objective was to determine if there is a “knee” in the sensitivity trend 

curve at a damping value less than 0.5%. 

Over the period from mid-December 2009 to early February 2010 the NASA Technical Fellow 

for GN&C, along with several members of his GN&C TDT, met multiple times with the MSFC 

Ares-I GN&C team. The following items were reviewed by the NESC assessment team: 

 Ares-I GN&C architecture, 

 Ares-I GN&C stability and flight performance requirements, 

 Legacy GN&C references from Saturn, 

 Analysis methodology and tools, 

 Analytic assumptions, 

 Uncertainty assumptions, 

 Modeling techniques, 
27 

 The technical interaction/model data transfer between the Orion team and Ares-I GN&C 

team, and 

 Analysis results, both frequency domain and time domain. 

The MSFC Ares-I GN&C Team has used technically sound and appropriate (traditional industry 

standard type) analytical methods to assess the impact of the very lightly damped (0.03%) Orion 

slosh mode. The depth of the MSFC team’s technical rigor, physical understanding, and problem 

insights was clearly apparent. 

Figure 14. Ares-I control, flexible body, and slosh mode frequencies vs. time. 

Specifically regarding the level of conservatism in the MSFC slosh analyses, the NESC 

assessment team concluded the following: 

 The Orion estimates of propellant slosh damping ratios in the Orion SM tanks of 0.03% 

(NTO) to 0.06% (MMH) are generally consistent with calculations using industry 

standard and historical techniques, and do not appear to be overly conservative. 

 The Orion-provided pendulum mechanical models for Orion SM propellant slosh, 

provided to the Ares-I personnel, and the integration and use of those models in the AresI MAVERIC 6-DOF simulation and linear analysis tools are generally consistent with 

industry standard and historical techniques, and do not appear to be overly conservative. 
28 

However, it was also observed that the MAVERIC 6-DOF simulation tool used by the Ares-I 

flight control engineers at MSFC does not clearly flag cases in which the displacement of the 

point mass used in the Orion SM tank slosh spring-mass mechanical model exceeds the 

maximum displacement consistent with the linear fluid slosh fundamental mode on which they 

are based. The NESC team felt strongly that this transition point, between the linear slosh regime 

and the non-linear slosh regime, must be clearly identified. 

For linear fluid slosh at the fundamental slosh mode to occur, there is a generally accepted rule 

of thumb that the maximum height of the fluid wave at the outer wall of the tank should not 

exceed approximately 10% of the radius of the tank. For greater fluid wave heights, the 

fundamental anti-symmetric slosh wave tends to break apart into more complex wave shapes 

because of splashing and breaking of the waves, reducing the amount of mass that continues to 

slosh at the fundamental mode natural frequency. Using the linear mechanical slosh model 

beyond this linear range could result in overly pessimistic predictions of the forces imparted on 

the tank walls by the propellant slosh. 

The Southwest Research Institute slosh tool computes an estimate of the ratio of the maximum 

fluid wave height at the tank outer wall to the maximum amplitude of the pendulum mechanical 

model. For the Orion propellant tanks at the nominal lunar and ISS mission fill levels, the slosh 

tool predicts that ratio to be approximately 1.5. That is, the maximum fluid wave height at the 

tank outer wall is approximately 50% larger than the maximum amplitude of the pendulum 

mechanical model. Since the Orion propellant tanks are 50 inches in diameter, 10% of their 

radius is equivalent to a maximum wave height at the tank outer wall of approximately 2.5 

inches, which when divided by the ratio of wave height to pendulum displacement, yields a 

maximum pendulum amplitude (or the equivalent spring-mass displacement) of approximately 

1.7 inches. 

Until the recent design cycle, the Ares-I MAVERIC 6-DOF simulation did not have the 

capability to output the time history of the Orion SM tank slosh spring-mass displacement. Only 

the statistics for maximum amplitude could be examined. However, prior to the analyses of the 

current design cycle, the simulation was modified to output time history data of the slosh model 

spring-mass displacement. Initial evaluations of the nominal and typical Monte Carlo time 

history data does appear to suggest that the Orion SM tank slosh spring-mass displacements may 

be beyond the maximum allowable range for linear slosh models. 

Based on this initial finding, the NESC team recommended that MSFC review all MAVERIC 

tank slosh displacement data (nominal as well as Monte Carlo dispersed) to confirm operation 

within the linear range of the spring-mass mechanical models. 

As mentioned above, the MAVERIC 6-DOF simulation used for the current Ares-I GN&C 

design cycle was modified to provide output time history data for the displacement of the springmass Orion tank slosh dynamics model. A comprehensive review of the tank slosh spring-mass 

displacement for nominal missions as well as Monte Carlo simulation ensembles should be 

performed to confirm whether the slosh model is operating within its linear range. Preliminary 

data provided for the slosh spring-mass displacement time histories has indicated that for low 

slosh damping, the Monte Carlo dispersed trajectory data does show significant intervals of 

operation of the Orion tank slosh dynamic model beyond its linear range. 
29 

Figure 15 shows a sample plot of the displacement of the spring-mass corresponding to the Orion 

NTO tank for a Monte Carlo run of mission TD71, using 0.1% tank slosh damping. In the figure, 

the red line indicates the nominal mission spring-mass displacement, while the blue lines indicate 

the 2000 Monte Carlo runs. The black horizontal line shows the approximate limit of 0.14 feet 

(or 1.7 inches) for operation of the slosh model in its linear range. Clearly, there are significant 

time intervals where the tank slosh displacement exceeds the maximum limit for linear operation 

for the dispersed cases, although the nominal simulation run shows operation well within the 

linear range. 

The GN&C TDT members of the assessment team recommended that a comprehensive review of 

existing MAVERIC simulation data should be reviewed to determine the extent at which the 

current Orion tank slosh spring-mass mechanical models may be operating outside of their linear 

range. Following such an analysis, the Ares-I personnel may consider developing and using nonlinear tank slosh dynamic models. Recent technical discussions of the NESC assessment team 

with representatives from United Launch Alliance have suggested that such non-linear tank slosh 

models have been developed for use with large tank slosh mechanical model displacements and 

have been used in their tank slosh analyses, which may be of use to the Ares-I analysis of Orion 

tank slosh. 

Therefore, it was also recommended that MSFC consider the development and use of non-linear 

tank slosh mechanical models for the Orion SM propellant tank slosh. The NESC team observed 

that the displacement of the Orion tank slosh spring-mass mechanical models may be operating 

beyond the linear range at which they are considered to be valid mechanical representations of 

Orion tank slosh dynamics. It was recommended that the Ares-I personnel consider 

implementing non-linear tank slosh mechanical models in their MAVERIC 6-DOF simulation 

and linear analysis tools to provide higher confidence in their predictions of the effects of Orion 

SM tank slosh on the Ares-I vehicle controllability. 

Figure 15. Orion upstream NTO tank slosh displacement (0.1% damping; TD71 mission). 

10% tank radius / 1.5 
30 

The primary recommendation of the NESC team was for CxP to modify Orion propellant tanks 

to increase damping. 

Any concerns raised about the “degree of confidence” in the Ares-I / Orion SM tank slosh 

dynamics models used for the analysis of the effects of Orion tank slosh on Ares-I controllability 

are largely related to uncertainties caused by the very low damping ratios predicted for the Orion 

SM tank slosh. For example, operating those slosh mechanical models within their linear range 

and the effects of lateral tank excitation near the fluid slosh fundamental mode natural frequency 

are largely affected by the low damping, and increasing that damping by an order of magnitude 

or more, would help improve the confidence in the applicability of those slosh mechanical 

models. 

4.6 Kepler Recurring Anomalies 

In January 2010, the NESC was requested to support an investigation into recurring anomalies 

onboard the Kepler spacecraft. This spacecraft, launched on March 7, 2009, is depicted in Figure 

16. The Kepler Mission (NASA Discovery mission #10) is specifically designed to survey our 

region of the Milky Way Galaxy to discover hundreds of Earth-size and smaller planets in or 

near the habitable zone and determine how many of the billions of stars in our galaxy have such 

planets. Results from this mission will allow us to place our solar system within the continuum 

of planetary systems in the Galaxy. 

At the time of this request, the spacecraft had been in routine science operations returning highquality science data for over nine months since operations began on May 12, 2009. The 

spacecraft had experienced a number of anomalies that have interrupted science data collection, 

but was operating well within its margins throughout the Science Operations mission phases and 

was still meeting its science requirements. However, while Kepler had not experienced any 

spacecraft emergencies it had experienced a number of star-tracker related anomalies. 

Figure 16. Kepler spacecraft and photometer. 
31 

The GN&C Technical Fellow and members of the GN&C TDT supported this NESC assessment 

of the Kepler anomalies. The NESC team reviewed the spacecraft’s general operations 

philosophy, current analysis of the anomalies, the timelines and fish-bones generated, the 

mitigations undertaken or under consideration, and the effects of the mitigations. The team also 

studied the mission’s anomaly response and resolution process as well as the available anomaly 

statistics. The NESC team also investigated the planned paths to get to the root-cause of the 

anomalies, the remaining work to bring anomalies to closure and any residual risks. 

The scientists using the Kepler data prefer long and stable data segments to perform their search 

for planetary transits. The mission design therefore requires only simple periodic operational 

interactions with the Kepler spacecraft. There are three key operational activities performed by 

the Kepler Flight Operations Team (FOT): 1) semi-weekly health and safety contacts; 2) 

monthly science data downloads via contact with NASA’s Deep Space Network (DSN); and 3) 

quarterly roll maneuvers. The Kepler photometer instrument stares at the same field for the entire 

mission. The science target location is redefined on the photometer’s charge coupled devices 

after each quarterly roll. There are also small sets of science target changes for Guest Observers 

and for higher-resolution observations. 

Kepler is a data-driven mission, so data completeness is key to mission success. There are two 

key data completeness requirements to be satisfied. The first is the 92% requirement for Onboard 

Data Completeness. This requirement essentially has to do with how much of total operational 

time is the spacecraft actually collecting science data. It is influenced by numerous factors, 

including the time spent off-pointing to downlink science data, the time needed to perform the 

quarterly rolls, the time needed to unload excess reaction wheel momentum and the time the 

vehicle spends in safe mode events or other anomalies. The second major requirement is the 

Flight-to-Ground Data Completeness metric of 99%, which is directly linked to how much of the 

science data collected onboard actually reaches the mission’s data archive. The quality of the 

DSN space-to-ground link performance and the ability for data retransmission primarily impacts 

this requirement. 

Since it so fundamentally drives the attainment of the Onboard Data Completeness requirement, 

the NESC team first considered the Kepler FOT team’s anomaly response philosophy. The 

response to a vehicle anomaly is guided by a realistic and pragmatic three-tiered priority 

philosophy. The first priority is to ensure vehicle safety by preserving functionality, including 

on-board fault protection. The second priority is to resume science data collection while the 

anomaly investigation proceeds in parallel. The third priority is to mitigate against further 

occurrences of the anomaly in order to reduce the threat of future loss of science or spacecraft 

functionality. The Kepler engineers budget for the potential loss of up to twelve days of science 

data collection per year from safing events. 

Since operations began on May 12, 2009, as of January 2010, the Kepler spacecraft has entered 

safe mode three times (June 15th , July 2nd , and November 18th 2009) and the loss of the 

finepoint attitude control capability needed for science data collection has also occurred three 

times (August 13th , September 14th , and October 9th 2009). Entry into safe mode is a preplanned self-protective measure that the spacecraft takes when something unexpected occurs 

onboard. During safe mode, the spacecraft points the solar panels directly at the sun and begins 

to slowly rotate about a sun-aligned axis. The spacecraft automatically powers off the 
32 

photometer and one redundant subsystem as a safety precaution. Shortly following the 

November 18th safe-mode entry the FOT engineers immediately began telemetry analysis to 

determine spacecraft subsystem health and root-cause determination of what triggered the safe 

mode. This was the third safe mode Kepler had experienced since operations began in May 2009 

and it was classed as a Criticality 1 anomaly. Criticality 1 anomalies are those that represent an 

unacceptable risk or have high impact to achieving the Kepler baseline mission objectives. 

Onboard detection of a Criticality 1 anomaly will automatically force the spacecraft into safe 

mode. 

This particular safe mode occurred when the team was preparing to download another month of 

scientific data from Kepler. The scientific data was not at risk and was subsequently downloaded 

successfully via the DSN communications link the day after the safe-mode event. Engineers 

verified nominal performance of all of Kepler’s systems and successfully recovered the vehicle 

from safe mode. Science data collection was resumed by the evening of November 20. The safehold entry was traced to an undesired interaction between the spacecraft’s flight software and its 

star tracker. The root cause is understood. Near-term mitigation of the problem is in place and a 

long-term mitigation is to be implemented. The anomaly investigation has revealed three startracker related issues. The first is that the star trackers unexpectedly drop stars and/or declare 

stars invalid, leading to a lost-in-space (LIS) condition. The second is that the star tracker 

remains in LIS mode for unexpectedly long time. The third and last issue is that star tracker 

directed search commands occasionally fail. 

The NESC team is still in the process of reviewing the details of the anomaly investigations and 

the associate mitigation or resolution plans. The Kepler management team, the NASA 

engineering team, and the contractor engineering team are using sound risk management 

principles to guide the process for the resolution of each anomaly and to set appropriate 

priorities. 

5.0 Conclusions 

This paper has described how the NESC was formed as an independent organization dedicated to 

promoting safety through engineering excellence. A resource for the Agency, it is a valuable 

problem-solving asset for the high-risk programs that NASA has always undertaken. NESC 

brings together some of NASA’s best engineers with experts from industry, academia, and other 

government agencies to address our highest risk, most complex issues. NESC strives to cultivate 

a safety-focused culture focused on engineering and technical excellence, while fostering an 

open environment and attacking the Agency’s technical challenges with unequalled tenacity. 

The NESC is more than a problem-solving organization however. It is also an organization that 

works to improve the competence of our entire engineering workforce through the opportunity to 

work on challenging problems, through exposure to other people, tools, techniques and facilities 

from across the Agency, through discipline advancing proactive work, and through its 

promulgation of lessons learned via technical reports and the NESC Academy courses. 

The backbone of the NESC is the ready group of engineering experts organized into 15 

discipline areas TDTs. In this paper, the purpose of the NESC GN&C TDT has been highlighted 

and a number of their experiences described. The members of the GN&C TDT have contributed 

to solving problems in many of NASA’s human spaceflight and robotic spaceflight programs and 
33 

projects. Their collective efforts have ranged from the assessment of alternative LASs, on-orbit 

robotic spacecraft anomalies, propellant sloshing impacts on launch vehicle flight stability and 

control, loads and dynamics support, and specialized ground testing of future flight systems. 

The NESC, because of its demonstrated ability to focus the technical talent from across all 

NASA Centers to bear on diverse high-priority problems, has become a valuable resource to 

senior Agency decision makers. The NESC has established itself as a reliable, credible and 

respected organization within the Agency and it is an outstanding example of Engineering 

Excellence in practice. This is evidenced by the increase in requests, from all levels of the 

Agency, for NESC support in resolving problems, reviewing activities, and conducting special 

studies. The NESC, employing the advantages of distributed “virtual organization” architecture, 

has shown its ability to concentrate appropriate levels of technical expertise efficiently when and 

where needed to independently address some of NASA’s most challenging and most visible 

problems. Some consider the NESC to be one of the Agency’s most positive post-Columbia

success stories. 

6.0 References 

1. Dennehy, C. J.: “NASA Engineering and Safety Center (NESC) GN&C Technical 

Discipline Team (TDT): Its Purpose, Practices and Experiences.” Presented at the AIAA 

Guidance, Navigation and Control Conference and Exhibit, Hilton Head, August 2007, 

paper number AIAA-2007-6332,. 

2. Abramson, H. N.: The Dynamic Behavior of Liquids In Moving Containers. NASA Space 

Vehicle Design Criteria Series, NASA Document SP-106, 1967. 

3. Anon.: Slosh Suppression. NASA Space Vehicle Design Criteria Series, NASA Document 

SP-8031, May 1969 
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