DOT/FAA/AR-04/19 

Office of Aviation Research 

Washington, D.C. 20591

Hazard Assessment for Usage 

Credits on Helicopters Using 

Health and Usage Monitoring 

System 

July 2004 

Final Report 

This document is available to the U.S. public 

through the National Technical Information 

Service (NTIS), Springfield, Virginia 22161. 

U.S. Department of Transportation 

Federal Aviation Administration
NOTICE 

This document is disseminated under the sponsorship of the U.S. 

Department of Transportation in the interest of information exchange. The 

United States Government assumes no liability for the contents or use 

thereof. The United States Government does not endorse products or 

manufacturers. Trade or manufacturer's names appear herein solely 

because they are considered essential to the objective of this report. This 

document does not constitute FAA certification policy. Consult your local 

FAA aircraft certification office as to its use. 

This report is available at the Federal Aviation Administration William J. 

Hughes Technical Center's Full-Text Technical Reports page: 

actlibrary.tc.faa.gov in Adobe Acrobat portable document format (PDF). 
 Technical Report Documentation Page 1. Report No.

DOT/FAA/AR-04/19

2. Government Accession No. 3. Recipient's Catalog No.

 4. Title and Subtitle

HAZARD ASSESSMENT FOR USAGE CREDITS ON HELICOPTERS USING 

5. Report Date 

July 2004

HEALTH AND USAGE MONITORING SYSTEM 6. Performing Organization Code

7. Author(s)

Mike Augustin

8. Performing Organization Report No. 

9. Performing Organization Name and Address

Bell Helicopter Textron Inc. 

10. Work Unit No. (TRAIS) 

 

Ft. Worth, TX 76101 11. Contract or Grant No.

 DTFA0397C00032 

12. Sponsoring Agency Name and Address

U.S. Department of Transportation 

Federal Aviation Administration

13. Type of Report and Period Covered

 Final Report

Office of Aviation Research 

Washington, DC 20591

14. Sponsoring Agency Code

 ASW-112

15. Supplementary Notes 

The FAA William J. Hughes Technical Center COTR was Dy Le.

16. Abstract 

A rotorcraft Health and Usage Monitoring System (HUMS) has two important roles. First, by accounting for maneuvers and 

conditions that are more severe than those that the aircraft was designed for, premature fatigue and other types of failures that 

could be catastrophic can be avoided. Second, by obtaining credit for service that is less demanding than those for which the 

aircraft was certified, longer service times can be justified, which will allow more economical operations. However, while usage 

monitoring has prevented failures, to date, no applicants have successfully followed the guidance of the Federal Aviation 

Administration Advisory Circular (AC) to develop individual aircraft usage credits. This report collects data and other 

information that was developed by Bell Helicopter Textron for its HUMS applications and provides it in a systematic manner to

support more general and informed use of this important technology by its incorporation into the AC. 

A detailed overview of the HUMS methodology is provided in this report that describes how flight condition recognition is used 

to determine an individual aircraft load spectrum that, in turn, is used to calculate the effective hours that correspond to this 

usage. The report then focuses on sources of error in usage monitoring and how these can be overcome. These include 

inaccurate data, missing data, and incorrect algorithms. Next, system compliance with usage applications is discussed. Finally, a 

prototype usage system for implementing and using HUMS is outlined, which has been evaluated for compliance with the HUMS 

AC. This analysis verified that the prototype system and its architecture are compliant with the intent of the AC, including

required parameter rates and accuracy levels. This report provides a valuable resource for usage monitoring that can both 

increase safety and enhance more economical operations, while also providing the basis for the eventual linkage of HUMS with

the rotorcraft damage tolerance methodology. 

17. Key Words 

Flight condition recognition, FCR; Rotorcraft damage 

tolerance, RCDT; Display panel; Sampling rates; Usage 

monitoring; Usage spectrum; Usage hazard assessment; Usage 

hazard analysis; Advisory Circular, AC 

18. Distribution Statement 

This document is available to the public through the National

Technical Information Service (NTIS) Springfield, Virginia 

22161.

19. Security Classif. (of this report)

 Unclassified

20. Security Classif. (of this page)

 Unclassified

21. No. of Pages 

 65 

22. Price

Form DOT F1700.7 (8-72) Reproduction of completed page authorized
TABLE OF CONTENTS 

Page 

EXECUTIVE SUMMARY vii 

1. INTRODUCTION AND OVERVIEW 1-1 

1.1 System Overview 1-1 

1.1.1 Airborne System Description 1-2 

1.1.2 Ground Station Description 1-3 

1.2 Usage Monitoring Overview 1-3 

1.2.1 Spectrum Damage 1-3 

1.2.2 Ground-Air-Ground Cycle Counting 1-11 

1.2.3 Parts Affected by Hoist Operation Cycles 1-11 

1.3 Other Nonrequired Functions 1-11 

1.4 Fault Tree Analysis 1-12 

2. INACCURATE DATA 2-1 

2.1 Incorrect Acquisition of Data by the Airborne Unit 2-1 

2.1.1 Sensor or Sensor Wiring Failure 2-1 

2.1.2 Sensor Rigged Incorrectly 2-1 

2.1.3 Integration of System Failure (Wiring and Power Supply) 2-1 

2.1.4 Malfunction of the HPU 2-2 

2.1.5 Momentary Interruptions 2-2 

2.2 Inaccurate Data Caused by Ground Processing 2-2 

2.2.1 Hardware Failure 2-2 

2.2.2 Software Failure 2-3 

2.2.3 Misprocessed Data 2-3 

2.3 Mitigating Factors for Inaccurate Data 2-3 

2.3.1 Cockpit Verification of Data Using the HUMS Display Panel 2-3 

2.3.2 Tape Backup System 2-4 

2.3.3 Usage System Integrity Check Using the Master Card 2-4 

2.3.4 The Release Report 2-4 

2.3.5 The Process Report 2-5 

iii
3. MISSING DATA 3-1 

3.1 Causes 3-1 

3.2 Detecting Missing Data 3-1 

3.3 Logbook Time Audit 3-1 

3.4 Gap Filling Missing Data 3-2 

4. WRONG ALGORITHMS 4-1 

4.1 Software Requirements Incorrect 4-1 

4.2 Software Design Incorrect 4-1 

4.3 Coding Errors 4-1 

4.4 Testing 4-1 

5. SYSTEM COMPLIANCE ASSESSMENT 5-1 

5.1 Sampling Rates 5-2 

5.2 System Accuracy 5-3 

5.3 The Effect of Gusts 5-4 

5.4 System Integrity and Redundancy 5-5 

6. USAGE METHODOLOGY 6-1 

6.1 Usage Process 6-2 

6.2 Release of Data 6-3 

6.3 Effective Hours Calculation 6-3 

6.4 Flight Condition Recognition Algorithms 6-4 

7. USAGE SYSTEM SOFTWARE CONSIDERATIONS 7-1 

7.1 Usage System Checks and Balances 7-1 

7.2 Usage System Operating System 7-1 

7.3 Usage System Source Code 7-1 

7.4 The FCR Algorithm Prototyping Using Stand Alone Software 7-1 

7.5 Local Database 7-1 

7.6 Level of Compliance With RTCA/DO-178B 7-2 

8. SUMMARY 8-1 

9. REFERENCES 9-1 

APPENDICES

A—412 HUMS Usage Parameter Accuracies

B—Error Analysis of Vh for the Bell Helicopter Model 412 

iv
LIST OF FIGURES 

Figure Page 

1-1 HUMS Block Diagram 1-2 

1-2 Component Life Determination Process Using Flight Condition Recognition 1-6 

1-3 Fault Tree Analysis For HUMS Usage Credits 1-13 

2-1 Example of Parameter Coupling During an Acceleration From Climb Airspeed to 

Vh Airspeed 2-7 

2-2 Example of Usage Rate Trending 2-8 

5-1 Model 412 Helicopter Mast Torque Response 5-2 

6-1 Flowchart of Data Flow in the Usage Ground Station 6-1 

6-2 Main FCR Decision Module 6-5 

LIST OF TABLES 

Table Page 

1-1 Components That Can Be Tracked With a Usage System 1-4 

1-2 Parameters Obtained From Airborne Unit 1-7 

1-3 Parameters Derived in Usage Ground Station 1-8 

1-4 Flight Conditions Recognized 1-9 

2-1 Reasonableness Limits 2-6 

5-1 Usage Parameters—Accuracies, Rates, and Sources 5-1 

5-2 Comparison of Usage Parameter Accuracy to Applied Spectrum Banding 5-3 

6-1 Normal Ranges of Parameters 6-4 

v
LIST OF ACRONYMS 

BHTI Bell Helicopter Textron Inc. 

BIT Built-in-test 

COTS Commercial off-the-shelf

DRU Data retrieval unit 

ETS Engine transition state 

FAA Federal Aviation Administration 

FCR Flight condition recognition 

GAG Ground-to-ground 

HA Hazard assessment 

HDP HUMS display panel 

HIRF High-intensity radiated field 

HPU HUMS processing unit 

HUMS Health and Usage Monitoring System

LLDC Low level DC 

NASA National Aeronautics and Space Administration 

OAT outside air temperature 

PAC Power assurance check 

PC Personal Computer 

POT Potentiometer 

RIN Retirement index number 

RSS Root sum squared 

RT&B Rotorcraft Track and Balance 

Vac never exceed velocity 

Vh Maximum horizontal velocity 

vi
EXECUTIVE SUMMARY

A rotorcraft Health and Usage Monitoring System (HUMS) has two important roles. First, by 

accounting for maneuvers and conditions that are more severe than those that the aircraft was 

designed for, premature fatigue and other types of failures that could be catastrophic can be 

avoided. Second, by obtaining credit for service that is less demanding than those for which the 

aircraft was certified, longer service times can be justified, which will allow more economical 

operations. However, while usage monitoring has prevented failures, to date, no applicants have 

successfully followed the guidance of the Federal Aviation Administration Advisory Circular 

(AC) to develop individual aircraft usage credits. This report collects data and other information 

that was developed by Bell Helicopter Textron for its HUMS applications and provides it in a 

systematic manner to support more general and informed use of this important technology by its 

incorporation into the AC. 

A detailed overview of the HUMS methodology is provided in this report that describes how 

flight condition recognition is used to determine an individual aircraft load spectrum that, in turn, 

is used to calculate the effective hours that correspond to this usage. The report then focuses on 

sources of error in usage monitoring and how these can be overcome. These include inaccurate 

data, missing data, and incorrect algorithms. Next, system compliance with usage applications is 

discussed. Finally, a prototype usage system for implementing and using HUMS is outlined, 

which has been evaluated for compliance with the HUMS AC. This analysis verified that the 

prototype system and its architecture are compliant with the intent of the AC, including required 

parameter rates and accuracy levels. This report provides a valuable resource for usage 

monitoring that can both increase safety and enhance more economical operations, while also 

providing the basis for the eventual linkage of HUMS with the rotorcraft damage tolerance 

methodology. 

vii/viii
1. INTRODUCTION AND OVERVIEW.

A rotorcraft Health and Usage Monitoring System (HUMS) has two important roles. First, by 

accounting for maneuvers and conditions that are more severe than those that the aircraft was 

designed for, premature fatigue and other types of failures that could be catastrophic can be 

avoided. Second, by obtaining credit for service that is less demanding than those for which the 

aircraft was certified, longer service times can be justified, which will allow more economical 

operations. However, while usage monitoring has prevented failures, to date, no applicants have 

successfully followed the guidance of the Federal Aviation Administration (FAA) Advisory 

Circular to develop individual aircraft usage credits. 

A previous study, funded by the FAA, NASA, and the U.S. Army Research laboratory, was 

conducted to assess the benefits of rotorcraft usage monitoring [1]. A separate report gives the 

operator’s perspective [2]. Other works providing background in this area have also been 

published including a recent paper at the 52nd American Helicopter Society Forum titled “Usage 

and Structural Life Monitoring with HUMS.” They adequately document the philosophy and 

rational that provide the basis of a Health and Usage Monitoring System (HUMS). 

This document presents an end-to-end hazard assessment (HA) prepared in accordance with 

reference 3 for the Bell Model 412 helicopter HUMS kit. The HA begins in section 1 with an 

overview of the system that is proposed to be installed on the M412 helicopter, and then details 

the fault tree analysis of the overall system, focusing primarily on the areas where certification 

credit is desired. Sections 2, 3, and 4 identify and discuss the contributing areas to overall 

system failure, including inaccurate data, missing data, and wrong algorithms. Section 5 

summarizes the proposed alternate means of computing flight hours and retirement index 

number (RIN) for purposes of the airworthiness part life limitations in Chapter 4 of the 412 

Maintenance Manual. The emphasis is placed on describing how the mitigating factors will be 

incorporated into the production system. Finally, section 6 reviews usage system software 

considerations and allocates the software level. 

1.1 SYSTEM OVERVIEW.

A basic assumption in this report is that the HUMS will be installed as an optional system that is 

not required for the operation of the aircraft. Vibration diagnostics and engine monitoring will 

be available but only as a nonrequired enhancement to the current approved maintenance 

practices. 

The functions that are generally performed by a HUMS fall into the following four categories. 

• Usage: Determine the incremental amount of component life used 

• Onboard Maintenance: Rotorcraft track and balance (RT&B) 

• Pilot Assistance: Logbook data and power assurance check (PAC) 

• Condition Monitoring (Health): Vibration diagnostics, exceedances, and structural 

overloads 

1-1
The data acquisition and processing system to perform these functions is presented schematically 

in figure 1-1. Certification credit is currently being sought only for that portion of this system

that is related to usage as shown in the figure. 

Aircraft

Sensors

Vibration

Sensors and

Blade

Tracker

Flight Data

Acquisition

Vibration Data

Acquisition &

Diagnostics

Data

Retrieval

Unit

FDR

PC

Memory

Card

Airborne System Flight Line Local Maintenance

HUMS

Display

Panel

Usage

Processing

Health

Processing

Blocks enclosed in dashed line are

Usage System Components

FIGURE 1-1. HUMS BLOCK DIAGRAM 

1.1.1 Airborne System Description.

The basic airborne system for usage monitoring generally comprises the following types of 

equipment. 

1. The HUMS processing unit (HPU). The HPU is a centralized data acquisition and 

processing unit that acquires data; converts analog, digital bus, and discrete inputs into 

digital form; preprocesses the data; provides display output to the HUMS display panel 

(HDP); provides continuous usage parameter data to a personal computer (PC) memory 

card interface for storage; and provides access to the data for a piece of ground support 

equipment called the data retrieval unit (DRU). The HPU performs a continuous built-intest (BIT) on the system elements. 

2. A set of sensors and transducers. These provide signals to the HPU through wiring 

harnesses. Many signals are provided to the HPU by connection to existing aircraft 

systems via harnesses added to the aircraft. 

1-2
3. The HUMS Display Panel. The HDP is mounted in the cockpit and displays operational 

information to the pilots. Internal to the HDP are two PCMCIA type II card slots. These 

card slots will accommodate FLASH memory cards, used to store continuous usage 

parameter data. The HDP will contain a BIT and will send the results to the HPU. 

1.1.2 Ground Station Description.

The ground station for a usage monitoring system generally consists of a commercial off-theshelf (COTS) PC with a tape backup system, a PCMCIA card slot, and a printer. 

Minimum system requirements will be published in the Configuration Index Document for the 

project. The only software being developed specifically for this project is the usage software. 

COTS support software will be used as discussed in section 7. 

1.2 USAGE MONITORING OVERVIEW.

This section presents an overview of the system to be used to determine the fatigue life usage for 

dynamic components. A list of the possible parts to be tracked by usage monitoring is given in 

table 1-1, which is derived from Bell Helicopter’s 412EP Maintenance Manual, Chapter 4. The 

actual parts for which usage monitoring will be performed will be a subset of the list given in the 

maintenance manual. Three different types of damage, and the basis by which usage monitoring 

will be used to determine the effects of damage on the listed parts, are: 

• Spectrum damage—based on actual aircraft spectrum

• Ground-air-ground (GAG) loading damage—based on RIN accumulation 

• Internal hoist operations—based on number of operations using the internal hoist in the 

right forward position 

These three types of damage are described in more detail in the following sections. Note that 

while all the parts included in table 1-1 are affected by spectrum damage, a few are also 

subjected to damage from GAG loading and hoist operations. These parts are delineated in 

table 1-1. 

1.2.1 Spectrum Damage.

Usage parameter data can be used to determine the time in each maneuver category (spectrum)

for each HUMS-equipped aircraft. This spectrum will be combined with existing fatigue 

strength and load level survey certification data to determine the amount of service life that has 

been expended on fatigue critical parts. Figure 1-2 presents a diagram of this methodology. 

1-3
TABLE 1-1. COMPONENTS THAT CAN BE TRACKED WITH A USAGE SYSTEM 

Part Nomenclature Part Number Airworthiness House 

Yoke Assembly 412-010-101-123 5,000 hrs (1)

Spindle Assembly 412-010-156-105 5,000 hrs 

Spindle Assembly 412-010-156-113 10,000 hrs 

Spindle Assembly 412-010-190-101 5,000 hrs 

Pitch Horn Assembly 412-010-149-105 5,000 hrs 

Pitch Horn Assembly 412-010-149-111 10,000 hrs 

Retention Bolt 412-010-124-105 5,000 hrs 

Retention Bolt (Expandable) 412-010-137-103 5,000 hrs 

Damper Bridge 412-010-104-101 5,000 hrs 

Damper Bridge 412-010-185-101 10,000 hrs 

Damper Bridge 412-010-185-109 15,000 hrs 

Damper Bridge 412-010-183-101 10,000 hrs 

Damper Bridge 412-010-183-109 15,000 hrs 

Damper Bridge 412-018-068-101 10,000 hrs/15 yr 

Damper Bridge 412-010-170-101 10,000 hrs/15 yr 

Fitting 412-010-111-101 5,000 hrs

Bracket Assembly, Pendulum Absorber 412-010-214-101 10,000 hrs 

Bracket Assembly, Pendulum Absorber 412-010-215-101 10,000 hrs 

Bracket Assembly, Pendulum Absorber 412-010-264-101 10,000 hrs 

Arm Assembly, Pendulum Absorber 412-010-217-101 5,000 hrs 

Arm Assembly, Pendulum Absorber 412-010-263-101 5,000 hrs 

Pitch Link Tube 412-010-425-113 5,000 hrs 

Pitch Link Bearing Assembly 412-010-182-101 5,000 hrs 

Pitch Link Rod End Bearing 412-010-412-101 5,000 hrs 

Pitch Link Rod End Bearing 412-010-438-101 5,000 hrs 

Pitch Link Rod End Bearing 412-010-400-101 5,000 hrs 

Swash Plate Link Rod End Bearing 412-010-412-101 5,000 hrs 

Swash Plate Link Rod End Bearing 412-010-448-101 5,000 hrs 

Swash Plate Link Rod End Bearing 412-010-400-103 5,000 hrs 

Swash Plate Link Tube 412-010-406-105 5,000 hrs 

Drive Hub Assembly 412-010-445-101 10,000 hrs 

Rephrasing Lever Assembly 412-010-403-109 1,250 hrs 

Rephrasing Lever Assembly 412-010-403-113 5,000 hrs 

Drive Link Assembly 412-010-405-101 5,000 hrs 

Drive Link Assembly 412-010-193-101 5,000 hrs 

Swash Plate Outer Ring 412-010-407-105 10,000 hrs 

1-4
TABLE 1-1. COMPONENTS THAT CAN BE TRACKED WITH A USAGE SYSTEM 

(Continued) 

Part Nomenclature Part Number Airworthiness House 

Swash Plate Support Assembly 412-010-443-101 5,000 hrs 

Swash Plate Support Assembly 412-010-409-105 5,000 hrs 

Swash Plate Support Assembly 412-010-453-101 5,000 hrs 

Gimbal Ring 412-010-404-101 9,000 hrs 

Collective Sleeve 412-010-408-003 9,000 hrs 

Collective Lever Assembly 412-010-408-101 10,000 hrs 

Main Rotor Mast Assembly 412-010-101-105 10,000 hrs(1) or 80,000 RIN(2)

Main Rotor Mast Assembly 412-010-101-121 10,000 hrs(1) or 80,000 RIN(2)

Cap, Retention 412-010-161-101 10,000 hrs 

Cap, Retention 412-010-171-101 10,000 hrs 

Cone 412-010-165-101 10,000 hrs

Cone 412-010-169-101 10,000 hrs

Drive Pin 412-010-166-101 10,000 hrs 

Upper Cone Seat 412-010-164-101 10,000 hrs 

Upper Cone Seat 412-010-174-101 10,000 hrs 

Upper Cone Seat 412-010-186-101 10,000 hrs 

Splined Plate Assembly 412-010-177-101 10,000 hrs 

Splined Plate Assembly 412-010-167-105 10,000 hrs 

Splined Plate Assembly 412-010-177-113 10,000 hrs(1) or 80,000 RIN(2)

Lower Cone Seat 412-010-178-101 10,000 hrs(1)

Lower Cone Seat 412-010-168-105 10,000 hrs(1)

Lower Cone Seat 412-010-056-105 10,000 hrs(1)

Cone 412-010-179-101 10,000 hrs

Adapter, Tail Rotor Drive Quill 212-040-206-103 5,000 hrs 

Gear-Spiral Bevel, Tail Rotor Drive Quill 212-040-151-101 5,000 hrs 

Planetary Spider 412-040-785-101 2,500 hrs 

Tail Rotor Drive Adapter, Coupling 412-040-634-101 5,000 hrs 

Tail Rotor Drive Adapter, Gearbox 412-040-625-101 5,000 hrs 

Tail Rotor Drive Adapter, Flanged 412-040-622-101 5,000 hrs 

(1) Four additional hours must be logged in for each hoist operation performed in the penalty center of gravity 

region (see figure 4-1 of reference 4 for a full explanation). 

(2) The measure that occurs first is the one that governs. 

A list of the measured onboard data parameters required for usage evaluations, together with 

their sources and recorded frequency rates, is presented in table 1-2. All the parameters in this 

list can be recorded in the aircraft on a removable memory card for postprocessing. 

Postprocessing will occur on the ground, using a PC and specially designed usage software. In 

1-5
addition to the parameters recorded in the airborne unit, a second group of parameters will be 

derived using measured parameters. These derived parameters are presented in table 1-3. 

EFFECTIVE HOURS

USED OF

COMPONENT LIFE

LOADS PER

COND.

COMPONENT

HOURS

USED

 COND.

or EVENT

N

S

FATIGUE

STRENGTH

FLIGHT LOAD

SURVEY

X

+

FLIGHT CONDITION TIME (min) SPECTRUM

CREATED BY

FCR

Retained From Original

Certification Data

FIGURE 1-2. COMPONENT LIFE DETERMINATION PROCESS USING FLIGHT 

CONDITION RECOGNITION 

To determine the actual operating spectrum, ground processing will use predefined algorithms to 

classify the flight conditions using the current values of the parameters presented in tables 1-2 

and 1-3. This process is referred to as flight condition recognition (FCR). These predefined 

algorithms have been formulated by analyzing data recorded during scripted flight profiles flown 

during the load-level survey, and in other model 412 scripted flights, as available. All flight 

conditions performed for the aircraft certification load-level survey [5] will be recognized. 

These conditions are presented in table 1-4. 

The FCR algorithms will be structured such that, if a condition cannot be identified, the time that 

the aircraft spent in that condition will be classified as unrecognized. Conservatively, the time 

accumulated as unrecognized will be counted towards the maneuver from the M412 load-level 

survey with the highest damage rate for each part. However, the time history data for all usage 

parameters for each instance of an unrecognized maneuver will be recorded in a unique file for 

subsequent analysis by Bell Helicopter Textron Inc. (BHTI). The time categories for each flight 

condition will be divided into three gross weight ranges, which correspond to those used in the 

load survey, i.e., less than 8,000 lbs, from 8,000 to 10,000 lbs, and from 10,000 to 11,900 lbs. 

1-6
TABLE 1-2. PARAMETERS OBTAINED FROM AIRBORNE UNIT 

No. Parameter Name Frequency Source

1 Pressure Altitude 1 Hz Air Data System (1)

2 Indicated Airspeed 1 Hz Air Data System (1)

3 Magnetic Heading 4 Hz Aircraft Instrument Sensor 

4 Pitch Attitude 4 Hz Aircraft Instrument Sensor 

5 Roll Attitude 4 Hz Aircraft Instrument Sensor 

6 Normal Acceleration (Nz) 8 Hz Added HUMS Sensor 

7 Main Rotor rpm 2 Hz Aircraft Instrument Sensor 

8 Engine No. 1 Torque 8 Hz Aircraft Instrument Sensor 

9 Engine No. 2 Torque 8 Hz Aircraft Instrument Sensor 

10 Mast Torque (where applicable) 8 Hz Aircraft Instrument Sensor 

11 Outside Air Temperature 1 Hz Air Data System (1)

12 Altitude Rate (vertical velocity) 4 Hz Air Data System (1)

13 Collective Stick Position 4 Hz 

14 F/A Cyclic Stick Position 4 Hz 

15 Lateral Cyclic Stick Position 4 Hz 

16 Tail Rotor Pedal Position 4 Hz 

Flight data recorder channel 

from dual-channel position 

sensor installed for F/A cycle 

stick 

17 Time and Date 0.25 Hz HUMS System

18 Pilot Entered Gross Weight Once per 

Flight (2)

Entered at Cockpit Display 

Panel (3)

19 On Ground Discrete 1 Hz HUMS System

20 Aircraft Flight Hours 1Hz HUMS System

(1) Dedicated air data computer for HUMS. Uses same pitot/static source as in the copilot’s display. 

(2) Fuel burn algorithm will decrement the gross weight as the aircraft is in flight. 

(3) If the pilot does not enter a value, the system defaults to the gross weight that causes the most damage for 

each component. The entry has to be reinitialized, or entered again, after each landing. 

1-7
TABLE 1-3. PARAMETERS DERIVED IN USAGE GROUND STATION 

Parameter Name Frequency Derivation Technique

Rate of Change Indicated Airspeed 1 Hz d/dt (Indicated Airspeed) 

Rate of Change of Magnetic Heading 4 Hz d/dt (Magnetic Heading) 

Pitch Rate 4 Hz d/dt (Pitch Attitude) 

Roll Rate 4 Hz d/dt (Roll Attitude) 

Moving Average of Rate of Climb 4 Hz Average of four previous samples, the current 

sample, and the next four samples 

Engine State 

1. Twin 

2. Single 

3. Auto 

4 Hz Rules When No Previous State is Available: 

(1) QE1 and QE2 > 10% = Twin; 

(2) QE1 and QE2 > 5% = Auto; 

(3) Either QE1 or QE2 > 5%, but not both = Single; 

Rules When Previous State is Available:

(1) Previous State = Auto 

a. QE1 and QE2 > 10% = Twin; ETS = A-> T 

b. QE1 and QE2 > 10% = Single; ETS = A-> S 

c. otherwise still Auto; ETS = None 

(2) Previous State = Single 

a. QE1 and QE2 > 10% = Twin; ETS = S-> T 

b. QE1 and QE2 > 5% = Auto; ETS = S-> A 

c. otherwise still Single; ETS = None 

(3) Previous State = Twin 

a. QE1 and QE2 > 5% = Auto; ETS = T-> A 

b. QE1 and QE2 > 5% = Single; ETS = T-> S 

c. otherwise still Twin; ETS = None 

Engine Transition State (ETS) 

1. Twin to Single: T->S 

2. Twin to Auto: T->A 

3. Single to Twin: S->T 

4. Single to Auto: S->A 

5. Auto to Twin: A->T 

6. Auto to Single: A->S 

7. None 

Rate of Change of Collective Stick 

Position 

4 Hz d/dt (Collective Stick Position) 

Rate of Change of F/A Cyclic Stick 

Position 

4 Hz d/dt (F/A Cyclic Stick Position) 

Rate of Change of Lateral Cyclic Stick 

Position 

4 Hz d/dt (Lateral Cyclic Stick Position) 

Combined Engine Torque 4 Hz QE1 + QE2

Peak and Valley (Pedal) 4 Hz True when the peak/valley cycle of the position 

parameter exceeds threshold within a defined time. 

Peak and Valley (Longitudinal 

Stick Position) 

4 Hz 

Peak and Valley (Lateral Stick 

Position) 

4 Hz 

1-8
TABLE 1-4. FLIGHT CONDITIONS RECOGNIZED 

Number Flight Condition

Time or 

Event FCR Module

1 Rotor Start Event On Ground 

2 On Ground Time On Ground 

3 Takeoff, flat surface (>3° side slope) Event Takeoff

4 Landing, flat surface (>3° side slope) Event Landing

5 Shutdown Event On Ground

6 Steady Hover at 314 rpm Time Low Airspeed 

7 Steady Hover at 324 rpm Time Low Airspeed 

8 Hover Right Turn Time Low Airspeed 

9 Hover Left Turn Time Low Airspeed 

10 Hover-Longitudinal Control Reversal Time Low Airspeed 

11 Hover-Lateral Control Reversal Time Low Airspeed 

12 Hover-Pedal Control Reversal Time Low Airspeed 

13 Right Sideward Flight Time Low Airspeed 

14 Left Sideward Flight Time Low Airspeed 

15 Rearward Flight Time Low Airspeed 

16 Normal Takeoff and Acceleration to Climb Airspeed Time Accel/Decel 

17 Normal Approach and Landing, Twin Engine Time Accel/Decel 

18 Normal Approach and Landing, Single Engine Time Accel/Decel 

19 Level Flight, Twin Engine—0.4 Vh at 314 rpm Time Level Flight 

20 Level Flight, Twin Engine—0.4 Vh at 324 rpm Time Level Flight 

21 Level Flight, Twin Engine—0.6 Vh at 314 rpm Time Level Flight 

22 Level Flight, Twin Engine—0.6 Vh at 324 rpm Time Level Flight 

23 Level Flight, Twin Engine—0.8 Vh at 314 rpm Time Level Flight 

24 Level Flight, Twin Engine—0.8 Vh at 324 rpm Time Level Flight 

25 Level Flight, Twin Engine—0.9 Vh at 314 rpm Time Level Flight 

26 Level Flight, Twin Engine—0.9 Vh at 324 rpm Time Level Flight 

27 Level Flight, Twin Engine—1.0 Vh at 314 rpm Time Level Flight 

28 Level Flight, Twin Engine—1.0 Vh at 324 rpm Time Level Flight 

29 Level Flight, Twin Engine—Vne at 314 rpm Time Level Flight 

30 Level Flight, Twin Engine—Vne at 324 rpm Time Level Flight 

31 Twin-Engine Full-Power Climb Time Climbs/Descents

32 Single-Engine Full-Power Climb Time Climbs/Descents

33 Low-Speed Cyclic Pull-Up Time Pull-Up 

34 High-Speed Cyclic Pull-Up Time Pull-Up 

35 Normal Acceleration From Climb Airspeed to 0.9 Vh Time Accel/Decel

36 Low Speed Right Turn Time Turn 

1-9
TABLE 1-4. FLIGHT CONDITIONS RECOGNIZED (Continued) 

Number Flight Condition 

Time or 

Event FCR Module

37 High-Speed Right Turn Event Turn 

38 Low-Speed Left Turn Time Turn 

39 High-Speed Left Turn Event Turn 

40 Level Flight Longitudinal Control Reversal Event Level Flight 

41 Level Flight Lateral Control Reversal Event Level Flight 

42 Level Flight Pedal Control Reversal Time Level Flight 

43 Deceleration From 0.9 Vh to Climb Airspeed Time Accel/Decel 

44 Twin-Engine Partial Power Descent Time Climbs/Descents 

45 Single-Engine Partial Power Descent Time Climbs/Descents 

46 Twin-Engine-Single-Engine Transition Full-Power 

Climb 

Event Power Transitions

47 Twin-Engine-Single-Engine Transition Level Flight Event Power Transitions 

48 Single-Engine-Twin-Engine Transition Partial Power 

Descent 

Event Power Transitions

49 Twin-Engine-Autorotation Transition at Low Speed Event Power Transitions 

50 Twin-Engine-Autorotation Transition at High Speed Event Power Transitions 

51 Autorotation Time Climbs/Descents

52 Twin-Engine Recovery From Autorotation Event Power Transitions 

53 Autorotation Right Turn Time Climbs/Descents 

54 Autorotation Left Turn Time Climbs/Descents 

55 Side-Sloped Landing ( ≥ 3° and < 5° side slope) Event Landing

56 Side-Sloped Takeoff ( ≥ 3° and < 5° side slope) Event Takeoff

57 Side-Sloped Landing ( ≥ 3° and < 10° side slope) Event Landing

58 Side-Sloped Takeoff ( ≥ 3° and < 10° side slope) Event Takeoff

59 Unrecognized Time All Submodules

Vh = Maximum horizontal velocity

Vne = Never exceed velocity 

The major benefit of usage monitoring that uses FCR is an accurate accounting of how the 

aircraft has been operated. This information may allow an increased time in service for 

components on aircraft that were operated less severely than the assumed spectrum used in 

certification calculations. To mitigate the effects of unanticipated maneuvers, any increase in 

service life shall be limited to a factor of two. Of even more importance, an increase in safety is 

achieved for those aircraft operated more severely than the assumed certification spectrum

because their parts will be retired from service earlier than the initial certification calculation 

would allow. 

1-10
The current recommended component retirement hours, which were established for use with 

actual aircraft hours, will be retained. HUMS effective hours will be determined for each 

component by HUMS and used in lieu of actual aircraft flight hours for the component life 

expended calculations. These calculations, in terms of effective hours, will replace actual 

aircraft hours only when the required aircraft parameter data are available and are valid. 

1.2.2 Ground-Air-Ground Cycle Counting.

The components on the M412 that are most susceptible to GAG cycles are also monitored using 

aircraft parameter data. Damage cycles are counted and summed in a cumulative index that is 

referred to as the RIN. The RIN for a new part is zero. The part is retired when a predetermined 

maximum allowable RIN number is attained. 

The RIN can be recorded manually or automatically with a torque event-monitoring algorithm.

The manual method consists of counting each takeoff or external lift cycle as one RIN. For 

logging operations, two RIN are counted for each external lift. The automated system measures 

either combined engine torque or measured mast torque and applies a rain-flow algorithm to 

determine the torque cycles incurred by the aircraft. Torque cycles below the endurance limit 

threshold are truncated because they do not cause damage to the components. Damaging torque 

cycles are processed using BHTI’s damage equation and the material constants determined by 

testing for the part. Incremental damage is then calculated and subsequently converted to an 

integer RIN value, which is used by the operator’s maintenance facility for determining when 

GAG-effected parts should be retired. 

1.2.3 Parts Affected by Hoist Operation Cycles.

Certain parts are adversely affected by operation when the internal hoist is installed in the right 

forward position. For these selected parts, 4 hours are logged against each part’s retirement time 

for each hoist operation. The pilot will still be required to enter these operations manually into 

the logbook to account for these hours. Logbook hoist cycles will be transferred to the usage 

ground station by a manual entry process. 

1.3 OTHER NONREQUIRED FUNCTIONS.

The following available functions are not required for usage part life credits: 

• The PAC processing is performed in the HPU, and both the data used in calculations and 

the calculated results are displayed on the HDP. 

• The vibration data are obtained from accelerometers specifically installed for HUMS. A 

blade tracker can also be installed when required for RT&B. The results are displayed on 

the RT&B page of the HDP. 

• The vibration analysis, RT&B, exceedance monitoring, and trend data are processed with 

a separate HUMS processor card. The results are downloaded by the DRU, which can 

provide a quick look at the results to identify possible special inspections or needed 

maintenance actions. 

1-11
• Data will be transported to a HUMS health monitoring ground station to store and further 

process data. This ground station also provides a more detailed look at the results. 

These nonrequired functions have no affect on the usage system. 

1.4 FAULT TREE ANALYSIS.

The fault tree presented in figure 1-3 and the analyses described in sections 2 and 3 focus on the 

usage monitoring aspect of the HUMS. The hazard assessment for the system installation is 

presented in reference 6. The fault tree presents a top down analysis of the usage monitoring 

system. The analysis starts with the worst-case condition, i.e., a fatigue life-limited part being 

left in service too long, which is considered a potential catastrophic failure condition. The 

analysis breaks down the potential faults that could be the cause(s) of this condition, and how 

each cause can be compensated for or prevented. 

Incorporating usage monitoring to determine when life-limited parts should be retired will only 

change one aspect of the original certification process: the spectrum that is used to determine 

part life. Since FCR will realize the actual spectrum each aircraft is flown, it is imperative that 

the spectrum is determined correctly. If the spectrum is not determined correctly, a life-limited 

part could be left on the aircraft too long, leading to the possibility that a dynamic component 

could fail during flight. 

The factors that can affect the correctness of the spectrum are: 

• Inaccurate data 

• Missing data 

• Wrong algorithms

These factors, along with mitigating factors that can be used to nullify the effects of each 

problem, are discussed in detail in the following sections. 

1-12
1-13/1-14

Dynamic System

Component

Failure

Actual Aircraft

Spectrum

Determined by

FCR Not Correct

Missing Data

(3.0)

Inaccurate

Processed Data

(2.0)

Lost PC Memory

Card

PC Memory Card

Not Installed

During flight

Card Capacity

Exceeded During

Flight

HUMS Inoperative

HUMS Not

Installed

Raw Data Aquired

in Airborne

System Incorrect

(2.1)

Incorrect

Processing on

Ground

(2.2)

Sensor or Sensor

Wiring Failure

Sensor Not

Rigged Correctly

Hardware Failure

(2.2.1)

Software Failure

(2.2.2)

PC Failure Disk Crash Usage Monitoring Program Failure Operating System Failure

Pilot Fails to

Report Flight

Hours

Failure to Perform

Parameter Audit

(2.3.1)

Backup System or

Procedural Faults

Primary System

Faults

Tape Backup

System Failure

(2.3.2)

Failure to Review

Process Report

(2.3.4.1)

Failure t

Tren

(2.3

Miss Processed

Data

(2.2.3)

Failure of

Independant

Verification

(2.3.3)

Wrong Algorithm

(4.0)

C

Failure to Test

Against

Requirements

(4.4)

Failure of

Independant

Verification

(2.3.3)

Logbook Audit Not

Performed

(3.3)

Momentary

interruptions in

Data; HIRF,

intermittent wiring.

Failure of Input

Parameter Validity

Checks

(2.3.4.1.1)

Logbook Audit Not

Performed

(3.3)

HPU Malfunction

Integration of

System Failure

FIGURE 1-3. FAULT TREE ANALYSIS FOR HUMS USAGE CREDITS 
2. INACCURATE DATA.

Inaccurate data can result from (1) the airborne system acquiring incorrect raw data or (2) the 

ground processing affecting the data in an adverse way. These topics are discussed in detail in 

this section. 

2.1 INCORRECT ACQUISITION OF DATA BY THE AIRBORNE UNIT.

Basic aircraft flight data are acquired from existing onboard sensors by the flight data portion of 

the HPU. The sensor inputs to the HPU are analog type, except for the air data, which are an 

ARINC 429 input. These data are digitized, formatted, and sent to the flight data recorder (FDR) 

(if installed) and to a removable memory card. The data from the memory card are transferred to 

the usage ground computer to perform component life calculations. Errors in this path can occur 

from the following, all of which are detectable: 

• Sensor or sensor wiring failure 

• Sensor not rigged correctly 

• Integration of system failure (wiring, power supply) 

• HPU malfunction 

• Momentary interruptions 

2.1.1 Sensor or Sensor Wiring Failure.

The failure of a sensor, or of the wiring between the HPU and a sensor, can be detected by 

validity signals and reasonableness limits that are provided by the HPU. Checks that will 

invalidate the data associated with failures of this nature will be incorporated into the groundbased software. A periodic parameter audit will also detect this class of failures. 

2.1.2 Sensor Rigged Incorrectly.

The maintenance manual requires that the transducers for collective, cyclic positions, and pedal 

positions be rerigged after the control system has been worked on. Parameter correlation checks, 

to be discussed in detail in section 2.3.5.3, would be used to detect misrigged transducers. Full 

throw control position checks after sensor maintenance will also detect a rigging problem. 

2.1.3 Integration of System Failure (Wiring and Power Supply).

A failure of the HPU due to broken power wiring or to power supply problems would cause 

either a total shutdown of the system or the flow of data to the PC memory card to be stopped. 

In either case, no data would be recorded. 

Because this situation would not introduce incorrect data, it is considered to be the same as a 

HUMS inoperative failure. These are covered in section 3. Intermittent wiring problems are 

addressed in the section 2.1.5. 

2-1
A failure from an output of the HPU to the PC memory card unit would cause a loss of data and 

would also be addressed in section 3. A wiring failure from the HPU to the HDP would prevent 

the display from working and would be detected during a parameter audit. 

2.1.4 Malfunction of the HPU.

A total failure of the HPU results in the PC memory card being devoid of parameter files. This 

failure is considered a HUMS inoperative failure and is covered in the section 3. A partial 

failure of the HPU would be detectable by BIT, by the usage software’s input parameter validity 

checks, or by cross checks by the pilot during a parameter audit. A malfunction of the HPU that 

causes one or more parameters to be inaccurate (no such failure conditions are known, but this 

analysis assumes there may be such a condition) will be detectable by the input parameter 

validity checks if the inaccuracy causes the parameter to go out of range. If an inaccurate 

parameter remained in range, it would be detectable by the periodic parameter audit. Any data 

taken during a period of time where parameter data may have been inaccurate will be 

invalidated. Invalid data are conservatively gap filled as discussed in section 4.4. Intermittent 

data and misprocessed data are covered in sections 3.1-3.3. 

2.1.5 Momentary Interruptions.

Momentary interruptions in the collection of data due to High Intensity Radiated Field (HIRF) or 

intermittent wiring will be detectable. Parameter validity checks are accomplished on the usage 

input data stream and these checks, along with parameter correlation, should catch intermittent 

problems with single parameters and will invalidate the input data for the duration of the 

problem. Invalid input data are conservatively gap filled. 

The HIRF susceptibility of the HPU will be qualified to 5 volts/meter. Testing will be performed 

at levels above 5 volts/meter to quantify the effects of HIRF on the data. If the test results show 

that HIRF could adversely affect the data in a way that cannot be detected, then the HIRF 

qualification level will be revised. I-HIRF effects are typically short lived, however, and are not 

expected to have any significant impact other than a brief loss of data. Such losses of data will 

be conservatively gap filled, which is discussed elsewhere in this document. 

2.2 INACCURATE DATA CAUSED BY GROUND PROCESSING.

Ground processing that corrupts accurate data obtained from the airborne unit may be the result 

of a software or hardware problem. 

2.2.1 Hardware Failure.

Hardware faults fall into three categories: 

1. Partial system failures such as disk crashes that could cause the loss of data. 

2. A full total equipment failure such as a PC failure. Because data loss resulting from a 

disk crash can be alleviated by keeping backup copies of all data, a tape backup unit is 

2-2
included in the ground station design. PC failure is recovered by replacing the PC, then 

reloading the data via tape backup. 

3. Intermittent equipment failures or processor glitches and is covered in section 2.2.3. 

2.2.2 Software Failure.

Software problems can be related to the operating system and to the usage program. Failure in 

the operating system can be minimized by using a robust COTS operating system. Windows NT 

has been selected for this purpose. However, should a fault occur, the result would be an aborted 

or locked up usage program. In this case, no process report would be available. However, an 

operator-initiated procedure requiring review of the process report (see section 2.3.5) will 

provide adequate opportunity to discover a failure of this nature. 

Problems with the usage software will be detectable either from error messages printed in the 

process report or from the fact that there will be no process report at all. A failure of the 

software that aborts the program will result in a missing process report, and the failure can, 

therefore, be discovered and corrected or worked around (e.g., by data gap fill). An algorithmic 

failure would lead to either excessive flight hours in the unrecognized flight regime category or 

abnormal component usage rates would be experienced. Appropriate error messages will be in 

the process report for these conditions. The nature of the error detection processing is discussed 

in section 2.3. 

Software problems will be reported to Bell Helicopter and then corrected using DO-178B 

guidelines for software change. The corrected software will then be provided to all usage ground 

station operators. 

2.2.3 Misprocessed Data.

Processor glitches or other failures causing misprocessed data that surface after initial testing 

would likely be detectable over a period of time by reviewing the trends of the data. Periodic 

end-to-end testing of the usage ground station software using a known set of inputs and outputs 

on the PC memory card will also be used to ensure that the usage system is not misprocessing 

data. The subject of this ground station end-to-end testing is covered in more detail in section 

2.3.3. 

2.3 MITIGATING FACTORS FOR INACCURATE DATA.

Incorrect data can be detected in several ways. First, the majority of the data that the HUMS unit 

provides for usage are also displayed to the pilot via his instruments. Second, the HUMS system

itself contains checks for out of range parameters. And, last, the software in the ground 

processing includes functions that can detect inaccurate data. 

2.3.1 Cockpit Verification of Data Using the HUMS Display Panel.

The same data required for usage are also sent to the HUMS display panel in the cockpit where 

the actual sensor parameter data for usage can be viewed on a utility page. Periodically, this 

2-3
display is used to compare the HPU output data to the sensor data, most of which can be read on 

the cockpit instruments. A portion of these checks can be performed on the ground while others, 

such as airspeed, may need to be performed in the air. This activity is called the parameter audit. 

The results of the parameter audit will be input into the usage ground station before any effective 

life calculations are released and used by the maintenance facility. The time interval on this 

audit will be flexible but is not expected to exceed 1 month. The primary motivation for the 

operator to accomplish this audit often is that should the audit fail, all data taken since the 

previous audit may have to be invalidated, and thus, the operator could loose all his usage credit 

for that time. 

Sensor errors will be apparent to the flight crew when their instruments do not function as 

required during flight, since HUMS receives most of its data from the same sources as the 

aircraft instruments, see specification in table 1-2. 

2.3.2 Tape Backup System.

A tape backup system is included in the ground station configuration to provide a way to 

preserve the hard drive data files against system failures. If data are lost due to a system failure 

or for any other reason and are not recovered, then the operator’s cost-effectiveness is penalized 

by the requirement to conservatively gap fill all the lost data. There is a profit incentive for tape 

backup to be designed into the system so that the operator will make frequent (possibly even 

daily) tape backups to reduce the amount of data lost in system failures. The operator will 

establish a procedure to ensure that a tape backup of the usage data is made frequently. 

To prevent the hard drive from becoming full, parameter data will be archived using the tape 

backup system. Parameter data will be archived after they are processed. 

2.3.3 Usage System Integrity Check Using the Master Card.

A special PC memory card will be prepared that contains known good parameter data for a 

fictitious test aircraft. The flight regimes recorded on this Master Card and the associated 

incremental part damage for the test aircraft will be known and documented in a system integrity 

check test procedure. The Master Card will be retained at the ground station and can be used at 

any time to check the integrity of the usage ground station. This can be done by examining the 

incremental damage to the test aircraft when the card is processed. No release of test aircraft 

data to the maintenance facility will be allowed in order to avoid confusion. 

At system setup, and at any time a change is made to the system, the usage system integrity 

check must be performed using the Master Card. Monthly use of the Master Card will be 

required to ensure that the usage ground station is functioning properly. This periodic system

integrity check constitutes an independent verification of the usage system since known inputs 

should produce known outputs. 

2.3.4 The Release Report.

There are two types of reports that are used to report the results given by the usage software. 

First, the Process Report documents the correct processing of each PC card. Second, the Release 

2-4
Report documents the results of system audits and the release of data to the maintenance facility. 

The Process Report is the more complex of the two and will be discussed in section 2.3.5. 

The Release Report documents the results of the usage calculations and the release of data to the 

maintenance facility. The operator’s procedures for releasing data to the maintenance data 

tracking facility requires that each Release Report be reviewed. An error in the software that 

aborts the program will result in a missing report that can be discovered and corrected, or 

worked around (data gap fill) when necessary. The Release Report provides confirmation that 

all applicable parameter data have been validated before the release of the usage calculation 

results based on those parameter data. 

2.3.5 The Process Report.

The Process Report documents the results from parameter data processing. It also documents the 

results of system error detection processing. The operator’s procedure for processing the 

parameter data requires review of the Process Report for each PC memory card. A failure of the 

software that aborts the program will result in a missing Process Report, and it can, thus, be 

discovered and corrected. A processing error could lead to either excessive flight hours in the 

unrecognized flight regime category or in abnormal component usage rates when compared to 

previous data history. These conditions will be noted on the Process Report. 

Excessive unrecognized data will be reported on the Process Report and will, thus, be detected 

by examination of the report. When this excess alert is reported, the operator will perform a 

parameter audit to make sure the data being collected are accurate. If this audit does not show 

any problems, then the operator will require the pilots and maintainers responsible for the 

aircraft to identify if any of their activities or actions may have affected the data. Possible causes 

could be changes in flying style, the mission, a replacement of parts, or any adjustments that 

have been made. If no changes are found that may have affected the data, the operator will 

contact Bell Product Support for assistance in resolving the occurrence of excessive 

unrecognized data. The normal Bell product support process will analyze the operator’s data and 

find the cause of the abnormality. 

Additional detail on this error detection processing, as well as other messages that may appear on 

the Process Report, are described in the following sections. While the periodic parameter audit 

is generally sufficient to determine the validity of the parameter data, some system functions 

have been added to the design in order to reduce the amount of data that are lost when a failure 

occurs. For example, if a sensor failure occurs one day after a monthly parameter audit that is 

not discovered until the next parameter audit, then a full month of data will have to be 

invalidated and gap filled. These added system functions are intended to detect faults and report 

them on the Process Report to prevent the loss of significant amounts of data. 

2.3.5.1 Parameter Validity Checks. 

Out of range parameters can be detected by validity bits or reasonableness limits. Table 2-1 

presents a list of the reasonableness limits for the M412 Helicopter. 

2-5
TABLE 2-1. REASONABLENESS LIMITS 

No. Parameter Limits

1 Pressure Altitude Greater than 27648 feet 

2 Airspeed Greater than 304 kts 

3 Pitch Attitude Greater than ±90° 

4 Roll Attitude Greater than ±90° 

5 Collective Position

6 Longitudinal Cyclic Position 

7 Lateral Cyclic Position 

8 Tail Rotor Pedal Position 

Outside of 1.5% to 98.5% of 

transducer range 

9 Normal CG Acceleration Outside of -3.2 g to +5 g 

10 Outside Air Temperature Outside of ±64°C 

11 Altitude Rate Outside of ±8192 ft/min 

Other parameters, such as the ARTNC 429 air data parameters, have validity bits associated with 

them. All synchro input parameters produce an abnormal periodic signal to alert the ground 

station that the excitation voltage for the synchro was not present. Synchro is a signal voltage 

indicating angular position as defined in ARINC STD 407. 

2.3.5.2 Tightly Bounded FCR Algorithms.

The algorithms used for FCR will be developed such that the predefined limits for the principal 

data parameters used for each maneuver definition will have lower and upper bounds. If the 

conditions of the data parameters do not match any of the maneuvers, then the condition is 

classified as unrecognized. 

An excessive amount of time spent in the unrecognized maneuver category is indicative of either 

a system problem or a maneuver that was not included in the original load-level survey. If an 

excessive amount of time in the unrecognized category is detected, then a parameter audit must 

be performed. If all usage parameters are found to be working properly, then the operator must 

have either performed a maneuver not in the load-level survey or there is an error in the 

algorithms for that maneuver. Continued excessive amount of time in the unrecognized category 

should be reported to Bell through normal product support channels. Bell will then review the 

time history data file to determine if the maneuver damage is covered by using the worst-case 

damage rate for each part on the aircraft. However, this will be an extremely improbable case, 

since maneuvers not included in the FCR recognition would have to be acrobatic in nature. 

2.3.5.3 Parameter Correlation Checks.

The parameter correlation module will detect mean shifts in data that could result from a 

misrigged sensor. Mean shifts will be detected by correlating multiple parameters. Also, usage 

software will use the rate of change of the parameter data, instead of the raw parameter data to 

minimize the effect of mean shifts. 

2-6
The parameter correlation module functions as follows. 

• Primary parameters used for FCR identification of a maneuver will be checked against 

other secondary parameters not used for identification. 

• The secondary parameter will be correlated to the primary parameter to check if the 

secondary parameter is in range. 

• All primary and secondary parameter correlation data will be determined by analyzing 

actual flight test data from the M412 aircraft. 

Figure 2-1 presents airspeed acceleration as an example of this kind of correlation. The rate of 

change of airspeed would be used to determine if the aircraft was accelerating and, at the same

time, the collective could be checked for a corresponding change. This is just one example of 

the many different parameter checks that will be implemented. 

FIGURE 2-1. EXAMPLE OF PARAMETER COUPLING DURING AN ACCELERATION 

FROM CLIMB AIRSPEED TO Vh AIRSPEED 

2.3.5.4 Usage Rate Check.

The test for abnormal component usage rates is based on a significant increase or decrease in the 

effective component life expended per flight hour. A database will be maintained within the 

usage ground station for each component part number that is tracked. This database will retain 

the effective component hours used per actual flight hour for each operation. If this rate of usage 

2-7
increases or decreases by a predefined factor, then the system operator would be prompted to 

determine if the mission profile had been significantly changed. Figure 2-2 presents an example 

of such a change. 

FIGURE 2-2. EXAMPLE OF USAGE RATE TRENDING 

If the actual mission profile had not changed, a full HUMS system parameter audit and system

integrity check would be required. If the checkout determined that incorrect data had been 

gathered by the HUMS, usage monitoring would revert to actual aircraft hours. This would be in 

lieu of FCR hours for the time period when the aircraft data were questionable. If the checkout 

found nothing wrong, then the data would be accepted as valid. 

Possible causes of excessive changes to the usage rate could result from changes in flying style 

or the mission, the replacement of parts, or adjustments. Not finding any changes that could 

have affected the data, the operator will contact Bell for assistance in resolving the occurrence of

abnormal component life consumption. Via the normal Bell product support process, the 

operator’s data will be analyzed and the cause of the abnormality found. 

2-8
3. MISSING DATA.

3.1 CAUSES. 

Missing data can be caused by a variety of reasons. The major ones include: 

• PC memory card failure or capacity exceeded during flight 

• PC memory card not installed prior to flight or lost subsequently 

• HUMS inoperative, malfunctioned or not installed 

For these and all other possible causations not included in the above list, the process described 

below will detect any missing data. 

3.2 DETECTING MISSING DATA.

Missing usage data can be detected by comparing the actual cumulative aircraft hours tracked by 

the airborne unit (parameter 20 in table 1-2) with the total time accumulated in the ground station 

by FCR usage processing. The cumulative aircraft hours maintained by the airborne unit are 

initialized from the logbook. If FCR time is less than actual aircraft hours, a time delta exists 

and usage data are missing. 

A mitigating procedure has been included in the design to reduce the criticality of the software 

accounting for aircraft hours. A procedural check of actual aircraft flight hours as recorded in 

the aircraft logbook against usage data hours is accomplished by the logbook time audit. 

3.3 LOGBOOK TIME AUDIT.

At a preset interval, the usage system administrator will enter the following data for the latest 

flight with the HUMS installed. These data are required so that the usage program can be certain 

that all of the actual aircraft hours have been accounted for. 

• Aircraft hours from the logbook 

• Date and landing time of last flight 

The usage software will interrogate the usage database and compute the total cumulative usage 

hours processed and then compare that number to the hours from the logbook. If the usage hours 

processed are not equal to or greater than the logbook hours, then there are missing data. The 

missing data will be replaced with gap fill if the operator cannot find them. 

If the usage hours exceed the actual aircraft hours by more than 2 percent, the process report will 

document this anomaly, but the effective hours for parts will not be reduced. This is because the 

logbook is considered the master record of flight hours and is correct by definition. 

Consequently, because the logbook hours are considered the authoritative source, component 

damage caused by excessive usage system hours are not removed from the damage 

accumulation. 

3-1
3.4 GAP FILLING MISSING DATA.

To correct, or gap fill, missing data, a conservative time delta will be added to the time 

accumulated for each part monitored by the usage system. The missing data will be filled with 

actual flight hours or historical usage data whichever is worse. The gap-filled data will be 

tagged so that if the memory card containing the missing data is found, the gap-filled time can be 

removed and replaced with the time calculated by the usage system. 

The system will also prompt the system operator to enter the number of flights conducted during 

this missing time. The system operator should also note any logging operation external lifts or 

hoist operations performed with the hoist installed in the right forward position in order to 

correct the RIN count (see sections 1.2.2 and 1.2.3 for more information). 

3-2
4. WRONG ALGORITHMS.

There are three potential causes for the usage system to be implemented with wrong algorithms. 

These are (1) the specified algorithms were not translated into software requirements correctly, 

(2) the software requirements were not translated into design correctly, and (3) the design was 

not translated into code correctly. Once the code has been implemented, the testing process will 

verify that all the capabilities specified in the software requirements documentation will function 

as specified. 

4.1 SOFTWARE REQUIREMENTS INCORRECT.

The software requirements will flow down from the System Requirements Document. 

Requirements traceability will be tracked through the development to ensure that all 

requirements are implemented. 

4.2 SOFTWARE DESIGN INCORRECT.

The top level design is reviewed against the software requirements during a preliminary design 

review. Later in the design process, a detailed design review is held with the systems engineers, 

fatigue engineers, quality assurance, and others to review the detailed design against the 

requirements. 

4.3 CODING ERRORS.

Individual programming team engineers will be responsible for coding the various units and 

modules of the software design. The team will then conduct a code walk-through of each portion 

of the design so that all members can contribute their individual experience to the overall design, 

and thus improve the software detail design and coding. 

4.4 TESTING.

Test cases will be chosen to demonstrate that all the software requirements have been properly 

implemented. Any test failures will be analyzed and a determination made whether they affect 

the end product or not. Code involved in test failures will be fixed if necessary and retested. 

Other code, which may be affected by the changes, will also be retested. Any code changed after 

the code has been baselined will be written up and tracked using a Bell standard software 

problem report. 

Flight testing will be undertaken to demonstrate that known flight regimes can be recognized by 

the usage software, and that the component damage is attributed accordingly. A flight test plan 

will be produced in which a scripted flight (or flights) will be conducted to provide known inputs 

for comparisons of the usage software results against the expected results. When the expected 

results are achieved, the algorithms will have been proven correct. 

4-1/4-2
5. SYSTEM COMPLIANCE ASSESSMENT.

This section addresses issues of compliance with the usage application and the HUMS Advisory 

Circular. In particular, the compliance of sampling rates, system accuracy, the effects of gusts 

on the methodology, and system integrity and redundancy are discussed. 

The list of required system parameters were shown in tables 1-2 and 1-3. For reference in this 

section, calculated root sum squared (RSS) accuracies, parameter rates, and sources are shown in 

table 5-1. Detailed calculations for the RSS parameter accuracies can be found in appendix A. 

TABLE 5-1. USAGE PARAMETERS—ACCURACIES, RATES, AND SOURCES 

No. Parameter Name

RSS 

System Accuracy Rate Source 

1 Pressure Altitude 40 ft 1 Hz Air Data System(1)

2 Indicated Airspeed 1 knot 1 Hz Air Data System(1)

3 Magnetic Heading 2.09° 4 Hz Aircraft Instrument Sensor 

4 Pitch Attitude 0.66° 4 Hz Aircraft Instrument Sensor 

5 Roll Attitude 0.66° 4 Hz Aircraft Instrument Sensor 

6 Normal Acceleration (Nz) 0.07 g /0.21 g max 

datum

8 Hz Added HUMS Sensor 

7 Main Rotor rpm 0.3% 2 Hz Aircraft Instrument Sensor 

8 Engine No. 1 Torque 1.13% 8 Hz Aircraft Instrument Sensor 

9 Engine No. 2 Torque 1.13% 8 Hz Aircraft Instrument Sensor 

10 Mast Torque (where applicable) 1.13% 8 Hz Aircraft Instrument Sensor 

11 Outside Air Temperature 2°C 1 Hz Air Data System(1)

12 Altitude Rate (vertical velocity) 40 feet/minute 4 Hz Air Data System(1)

13 Collective Stick Position 3% 4 Hz HUMS System

14 F/A Cyclic Stick Position 3% 4 Hz HUMS System

15 Lateral Cyclic Stick Position 3% 4 Hz HUMS System

16 Tail Rotor Pedal Position 3% 4 Hz HUMS System

17 Time and Date System Setup 0.25 Hz HUMS System

18 Pilot-Entered Gross Weight 50 lbs (Pilot Entry) Once Per 

Flight(2)

Entered at Cockpit Display

Panel(3)

19 On Ground Discrete N/A 1 Hz HUMS System

20 Aircraft Flight Hours Audit to Pilot 

Logbook 

1 Hz HUMS System

(1) Dedicated air data computer for HUMS. Uses same pitot/static source as copilots display. 

(2) The fuel burn algorithm will decrement gross weight as aircraft is in flight. 

(3) If Pilot does not enter value, the system shall default to the gross weight causing the most damage for each 

component. Entry shall be reinitialized or entered again after each landing. 

5-1
5.1 SAMPLING RATES.

The sampling rates that were chosen for each parameter were based on sensor signal 

characteristics and the software update rate. In current practice, flight time is accumulated by 

the pilot and is logged to the nearest minute. The Bell Usage Credit System logs flight regime

data to a resolution of 1 second. The parameter acquisition rates were determined by assessing 

the response characteristics of the required sensor signals to dynamic changes in the flight 

regime responses. The minimum sample rate was therefore chosen to be 1 second for aircraft 

state parameters such as pressure altitude, airspeed, outside air temperature (OAT), and 

air/ground state discretes. Control position inputs, aircraft attitude, and heading rates were set at 

4 Hz to allow the software to resolve and anticipate the flight regime changes required to meet 

the overall system resolution of 1 Hz. Engine torques, mast torque, and Nz, all of which have a 

direct effect on component damage are sampled at a rate of 8 Hz to provide a correct digital 

representation of the respective signals. Figure 5-1 shows that an 8-Hz sample rate is sufficient 

to capture a good digital representation of the torque signal. 

FIGURE 5-1. MODEL 412 HELICOPTER MAST TORQUE RESPONSE 

5-2
5.2 SYSTEM ACCURACY.

The underlying accuracy of the HPU analogue acquisition is 0.2%. This is generally 5 to 10 

times more accurate than the aircraft instruments. The true root mean square accuracies of the 

entire system, including the sensors, were given in table 5-1. For usage analysis, the basic 

requirement of the system is that the accuracy be much greater (e.g., 10X) than the bands that are 

used in the certification spectrum of the aircraft. This is very conservative since in Bell’s 

proposed implementation of usage, all regimes are bumped up to the next level of the 

certification data. Assuming a uniform data distribution, the 10X accuracy goal means that 90% 

of the time the spectrum data will be put into the next higher band. For the remaining data 

points, the spectrum will be recorded in the existing band for a low-tolerance parameter. The 

basic design goal is to compute the individual aircraft spectrum into approximately 60 flight 

regimes with three gross weight bands for each flight regime in the certification load-level 

survey. 

Table 5-2 presents a comparison of the calculated RSS accuracies of the usage parameters to the 

corresponding banding as applied to the regime definitions found in table 1-2. Banding is 

applicable to pressure altitude, airspeed, and gross weight. It is indirectly applicable to some

additional calculations such as fuel burn corrections and calculations of Vh. Where banding is 

not applicable, the full-scale parameter accuracy is substituted. 

TABLE 5-2. COMPARISON OF USAGE PARAMETER ACCURACY TO APPLIED 

SPECTRUM BANDING 

No. Parameter Name

RSS 

System Accuracy

Applied 

Banding 

Accuracy as a Percent 

of Banding 

1 Pressure Altitude 40 ft 3000 ft 1.3% 

2 Indicated Airspeed 1 knot 12 knots 8.3%

3 Magnetic Heading 2.09° 360° 0.6%

4 Pitch Attitude 0.66° 90° 0.7%

5 Roll Attitude 0.66° 90° 0.7%

6 Normal Acceleration (Nz) 0.07g, 0.21 g max 

datum

-3 g, +6 g 0.8%, 2.3%

7 Main Rotor rpm 0.3% (1 rpm) 10 rpm 10%

8 Engine No. 1 Torque 1.13% 100% 1.13%

9 Engine No. 2 Torque 1.13% 100% 1.13%

10 Mast Torque

(where applicable) 

1.13% 100% 1.13%

11 Outside Air Temperature 2°C N/A Used to Calc Density Alt 

12 Altitude Rate (vertical 

velocity) 

40 feet/minute 600 feet/minute 6.7%

13 Collective Stick Position 3% 100% Used to Validate FCR 

5-3
TABLE 5-2. COMPARISON OF USAGE PARAMETER ACCURACY TO APPLIED 

SPECTRUM BANDING (Continued) 

No. Parameter Name

RSS 

System Accuracy

Applied 

Banding 

Accuracy as a Percent 

of Banding 

14 F/A Cyclic Stick Position 3% 100% Used to Validate FCR 

15 Lateral Cyclic Stick Position 3% 100% Used to Validate FCR 

16 Tail Rotor Pedal Position 3% 100% Used to Validate FCR 

17 Time and Date System Setup N/A See item 20 

18 Pilot-Entered Gross Weight 50 lbs (Pilot Entry) 2000 lbs 2.5%

19 On Ground Discrete N/A N/A 

20 Aircraft Flight Hours Audit to Pilot 

Logbook 

N/A Logbook Audit

Additional parameters are derived as referenced in table 1-3. These parameters are of four types: 

(1) the rates of change in airspeed, heading, pitch, and roll that are used for validating the 

transitions to various flight conditions, (2) engine transition states based on simple engine torque 

logic, (3) peak and valley algorithms for the controls, and (4) two additional derived parameters 

that are of importance, the initial gross weight and Vh. It should be noted that 

• Initial gross weight is corrected for fuel burn based on OAT, engine torques, and pressure 

altitude. This is a conservative calculation that is found in the Bell 412 Flight Manual. 

• Vh is calculated as a function of altitude and gross weight. The original spectrum is 

divided into conditions of 0.4, 0.6, 0.8, 0.9, and 1.0 Vh. Vh is a function of density 

altitude (pressure altitude and OAT) and gross weight. Vh is originally determined by 

measurement during the load level survey for the original usage spectrum. A worst-case 

analysis was made to determine the worst-case accuracy of Vh based on variations in 

pressure altitude, OAT, and gross weight. The results were good with a worst-case 

accuracy of less than 2% up to 12,000 ft, and less than 3% error at 15,000 ft. For most of 

these regions, the worst-case errors were actually less than 0.5%. The detailed error 

analysis for Vh can be found in appendix B. 

It should further be noted that the nominal value of Vh that is calculated with this expression is 

approximately 6 knots below the Flight Manual Values. This adds additional conservatism to 

the usage calculations. 

5.3 THE EFFECT OF GUSTS.

Wind gusts do affect the loads on helicopter dynamic components. This effect is accounted for 

in two ways. First, the load-level survey is flown in real conditions. While no effort is made to 

specifically seek turbulent air, there is also no effort to avoid it unless it directly conflicts with 

5-4
hitting a target, e.g., a very specific airspeed and second by performing control reversals. 

Control reversals, which are designed to simulate a pilot’s response to wind gusts, consist of 

rapid oscillations of the fore and aft, lateral, and pedals, each of which is performed individually. 

Control reversals are typically flown at both high speed and hover. This approach to handling 

the effect of gusts is part of the original load-level analysis and is not altered by the application 

of the current methodology for usage. 

5.4 SYSTEM INTEGRITY AND REDUNDANCY.

System integrity and redundancy, which are covered in sections 2, 3, and 4, include three areas 

of concern: inaccurate data, missing data, and wrong algorithms. The proposed usage 

methodology addresses all of these possible integrity issues without the need for any hardware or 

software redundancy. 

5-5/5-6
6. USAGE METHODOLOGY.

This section contains an outline of the end-to-end process being proposed for usage monitoring, 

including airborne operation and all aspects of ground-based processing. A chart showing the 

data flow for the portions of this process that are carried out in the ground station is presented in 

figure 6-1. Note that the maintenance data tracking facility responsibilities are not part of the 

Bell usage system but are functions of the operator’s existing system. 

Raw Parameter Data

from Aircraft via the

PC Memory Card

Copy PC memory

Card to Hard Drive

File on Usage Ground

Station PC

Update

Processing Log

File

Verify Copy

Reformat PC Memory

Card for Next Use

Update

Processing Log

File

Apply Proper

Engineering Unit

Conversions

Run FCR Algorithms

Run appropriate parameter

correlation algorithm as each

maneuver is determined

Save the Interval

Spectrum(s)

Produced for Current

PC Memory Card

Calculate Effective

Life Used for Each

Component

Gap Fill

for

Missing Data

List Data Parameters

that Do Not Correlate

on Process Report

Create Effective Life

File For Components

Card Download Operation

Process Data from Card Check for

Unreasonable,

Invalid, or Missing

Data

Provide results on the

Processing Report

Update Processing

Log File

Verify that all aircraft

flight hours are

accounted for

Produce the

Release Report

Produce the

Process Report

Update Processing

Log File

Verify that all aircraft

sensors were good

during the period

Release Data To Maintenance Facility

FIGURE 6-1. FLOWCHART OF DATA FLOW IN THE USAGE GROUND STATION 

6-1
Also included in this section is a discussion of some of the more important aspects of the 

methodology. Policies and procedures that an operator must comply with in order to utilize 

usage credits will be contained in the Usage System Operation Manual. These policies and 

procedures will expand upon all user actions mentioned in this report; in particular, the 

methodology that is presented in the following sections. 

6.1 USAGE PROCESS.

The end-to-end process is as follows: 

a. Install the PC memory card in the HUMS-equipped aircraft 

b. Conduct aircraft operations 

c. Remove the PC memory card and deliver it to the usage ground station 

d. Perform the card download operation by: 

• copying the PC memory card data to the usage hard drive. 

• confirming the data copy and preparing the PC memory card for its next use. 

e. Process data from the memory card by: 

• checking for unreasonable, invalid, or missing data; apply engineering units. 

• running the FCR algorithms and correlate the appropriate parameters for each 

maneuver identified. 

• creating an interval spectrum for each operation, and save the accumulated 

interval spectrums until the next data release. 

• Computing the life expended for all operations on this PC card by trending the 

usage rate. If there are unacceptable shifts in the trends, require interviews with 

pilot(s) and maintainer(s) about past missions versus present missions. If there 

was no mission or aircraft change, then require a full parameter audit. 

• producing a Process Report for the PC memory card data processed in which the 

anomalies and actions that should be taken are listed, including the parameters 

that should be audited from the parameter correlation checks. 

• checking for excessive time in the unrecognized flight regime, and requesting a 

full system parameter audit, if required, and developing a reminder message for 

monthly audits, checks, or system calibrations. 

f. If a parameter audit was performed, have user input the results 

g. If a system integrity check was performed, have user input the results 

h. When an aircraft logbook audit is performed, make sure all flight hours are accounted for 

i. Release data to a maintenance data-tracking facility in which: 

6-2
• gap fill for missing data (including where data was found invalid or unreasonable) 

will be made. 

• a Release Report will be produced in which each operation processed will include 

the operation time (flight time for each operation), the number of flights, and the 

damage to the parts for each operation. 

The maintenance data-tracking facility, which is not part of the Bell usage system, (1) receives 

component usage data from the ground station for a release cycle, (2) accumulates component 

usage over the life of the component, (3) maintains aircraft configuration, and (4) reports to the 

maintainer when a part should be replaced. 

6.2 RELEASE OF DATA.

When the system administrator wishes to prepare a report of effective incremental damage hours 

for all the components of a particular aircraft, then the following list of items will be checked 

internally by the usage software. 

• Can the logbook time be verified with the processed time?

• Are all the gaps in the data filled? 

• Has the parameter audit been completed? 

• Have all anomalies reported by process reports been explained or corrected?

Effective component hours will be considered valid for only the operations through the point in 

time at which (1) the last parameter audit was completed, (2) the last logbook time verification 

was completed, and (3) all anomalies were explained or corrected. A Release Report will then 

be issued. 

The Release Report will have two forms, one suitable for review and audit by the system

administrator, and one suitable for transfer to the operator’s maintenance data-tracking facility. 

6.3 EFFECTIVE HOURS CALCULATION.

Effective component hours are defined as the component life expended in accordance with 

HUMS usage algorithms. This value will be expressed in terms such that the original retirement 

life of the part established by the manufacturer will still be used. This methodology will allow 

operators to retain their current systems for tracking cumulative hours used on components, and 

allow components from a non-HUMS aircraft to be moved to a HUMS aircraft and vice versa. 

Aircraft conditions will be tracked by the number of events that have occurred or by time in a 

flight condition. The total effective component hours can be expressed in general as 

Total Effective Component Hours = (Effective Component Hours)Time +( Effective Component Hours)Event

= ∑ (Time in Flight Condition x Component Hours Used per flight condition)

+ ∑ (Number of Events x Component Hours Used per Event)

In these relations, time-in-flight condition denotes the total minutes that are spent in each flight 

condition, component hours used per flight condition is the equivalent hours of a given 

6-3
component’s life used per minute in that flight condition, number of events is the total number of 

given events performed during an operation, and component hours used per event is the 

equivalent hours of a given component’s life used per event. 

In most cases, the value of the effective hours is based either on a time-based or an event-based 

calculation. In a very few cases, these two methods are combined, as shown in the above 

equation. However, most often the individual terms of this equation are used to calculate either a 

time-based or an event-based number of effective hours. 

As an illustrative example, consider a flight that included 10 minutes of flight condition A for 

which the component hours established during the original flight loads survey of the aircraft for 

part X would be 0.7. Then, the effective component hours for that part in that flight would 

simply be (10/60) x 0.7 = 0.117 hrs. This process would then be repeated for each condition, 

with the results summed to obtain the effective component hours for part X. 

There will be a unique component hours used per flight condition or component hours used per 

event value for each component and flight condition combination. These values will be 

calculated using the same certification flight load-level survey data [5] and certification fatigue 

strength data [7] for the model 412 aircraft. 

6.4 FLIGHT CONDITION RECOGNITION ALGORITHMS.

Algorithms for FCR will be designed on a decision tree approach, with most decisions involving 

the test of either a measured or a derived parameter against a normal range. The normal ranges 

for the parameters have been established by the value of each parameter in steady-state 

conditions of level flight or hover. The normal ranges for the parameters are presented in 

table 6-1. An example of what may be used for the main decision module is presented in 

figure 6-2. Once a submodule has been entered, appropriate aircraft parameters, or combinations 

of parameters, are tested to determine if the aircraft is in a climb, dive, left or right turn, pull-up, 

etc. 

TABLE 6-1. NORMAL RANGES OF PARAMETERS 

Parameter High Value Low Value 

Pitch Attitude 15 deg -15 deg 

Pitch Rate +3.5 deg/sec -2.0 deg/sec 

Roll Attitude +4 deg -4 deg 

Roll Rate +4 deg/sec -4 deg/sec 

Normal Acceleration (Nz) +1.3 g’s +0.8 g 

Main Rotor RPM 339 rpm 300 rpm

Altitude Rate (vertical velocity) 600 ft/min -600 ft/min 

Collective Stick Position 105% 15% 

F/A Cyclic Stick Position 20% -80% 

Lateral Cyclic Stick Position 50% -50% 

Tail Rotor Pedal Position 40% -40% 

Rate of Change Magnetic Heading +4 deg/sec -4 deg/sec 

6-4
Yes

If On Ground (current) = True

And On Ground (previous) =

True

On Ground

If On Ground (current) = False

And On Ground (previous) =

True

Yes Takeoff Module

No

If On Ground (current) = True

And On Ground (previous) =

False

No

Yes Landing Module

Airspeed < 20 Kts

No

Low Airspeed

Module Yes

Pitch Rate Not Normal

No

Yes Pullup Module

Roll Rate Not Normal

No

Yes Turn Module

ROC Airspeed Not Normal

No

Accel/Decel

Module Yes

Vertical Velocity Not Normal

No

Climbs/Descent

Module Yes

Level Flight Module

No

Start Peak and Valley

Processing

For Lateral,

Longitudinal, and

Pedal Positions.

Required to detect

control reversals

Run

 Concurrently

ETS ≠ None Engine Transition

Module Yes

No

FIGURE 6-2. MAIN FCR DECISION MODULE 

6-5/6-6
7. USAGE SYSTEM SOFTWARE CONSIDERATIONS.

7.1 USAGE SYSTEM CHECKS AND BALANCES. 

In order to lower the usage software’s contribution to the overall system failure probability, 

monitoring algorithms will be implemented to crosscheck the usage algorithms. These include 

usage rate check, parameter correlation checks, and parameter validity checks. These algorithms

are discussed in section 2.3.5, Process Report.

Procedures have been included in the system design to further lower the usage software’s 

contribution to the overall system failure probability. These include review of the process report 

(see section 2.3.5), review of the release report (see section 2.3.4), the parameter audit (see 

section 2.3.l), the logbook audit (see section 3.3), the tape backup (see section 2.3.2), and the 

usage system integrity check (see section 2.3.3). 

7.2 USAGE SYSTEM OPERATING SYSTEM.

The operating system for the usage monitoring system ground station will be Windows NT, as 

suggested in section 2.2.2. 

7.3 USAGE SYSTEM SOURCE CODE.

The source code for the usage system software will be written in a high-level language. Highlevel languages typically provide for strong type checking and structured language constructs 

that provide a more maintainable end product and with reduced errors during development. 

7.4 THE FCR ALGORITHM PROTOTYPING USING STAND ALONE SOFTWARE.

All usage algorithms will be prototyped and then tested using existing flight test data from an 

aircraft instrumented to record parameter data required for FCR. During these flight test data 

flights, the maneuvers performed should be documented in the pilot’s log. 

7.5 LOCAL DATABASE.

The Bell usage credit system includes two databases and a configuration change file. A 

spectrum database will be maintained for each aircraft that is monitored with a usage system. 

The data will consist of (1) the cumulative spectrum, (2) effective hours for each operation and 

part combination for all operational time since the two previous logbook and parameter audits, 

(3) all files which contain parameter data for instances of unrecognized data, and (4) the usage 

rate per component for each aircraft. 

An aircraft configuration database will be included within the system. The configuration 

database will contain all aircraft in the usage monitoring program along with the part numbers 

and serial numbers of all usage components. This database will contain actual aircraft hours, 

effective hours, and total hours gap-filled for all the aircraft and components being monitored. 

7-1 
In addition to the configuration database, an electronic log file will be maintained of all 

component changes made on each aircraft. The file will include the aircraft serial number, 

change date, change time, component part numbers and serial numbers for the removed and 

added components, component disposition, and aircraft flight hours at the time of the change. 

7.6 LEVEL OF COMPLIANCE WITH RTCA/DO-178B.

The usage application software will be developed and tested to the guidelines of the HUMS 

Advisory Circular and RTCA/DO-178B. A usage credit application falls into the 1309 criticality 

category of hazardous/severe/major. 

Mitigation of the airborne software criticality is achieved through initial system calibration and 

periodic parameter audits. In the proposed application, this mitigation reduces the required 

airborne software level for flight parameter acquisition from Level B to Level C. 

The HUMS Advisory Circular does not allow a similar mitigation for the ground station

application software due to the use of COTS hardware and software. Therefore, the usage 

application software must be developed in accordance with RTCA/DO-178B, Level B. This 

hazard assessment shows the software contribution of the usage application software to meet or 

exceed the guidelines of the Advisory Circular. 

The integrity of the COTS software and hardware is assured (per the HUMS Advisory Circular) 

through use of an independent means of verification during the introduction to service period. 

7-2 
8. SUMMARY.

This report collects data and other information that have been developed by Bell Helicopter 

Textron for its Health and Usage Monitoring System (HUMS) applications, and provides it in a 

systematic manner to support more general and informed use of this important technology. A 

detailed overview of the HUMS methodology is provided first in this report that describes how 

flight condition recognition (FCR) is used to determine an individual aircraft load spectrum that, 

in turn, is used to calculate the effective hours that correspond to this usage. The report then 

focuses on sources of error in usage monitoring and how these can be overcome. These include 

inaccurate data, missing data and incorrect algorithms. Next, system compliance with usage 

applications is discussed. Finally, a prototype usage system for implementing and using HUMS 

is outlined. The prototype usage system has been evaluated for compliance with the HUMS 

Advisory Circular. This analysis verified that the prototype system and its architecture are 

compliant with the intent of the advisory circular, including required parameter rates and 

accuracy levels. This report provides a valuable resource for usage monitoring that can both 

increase safety and enhance more economical operations, while also providing the basis for the 

eventual linkage of HUMS with the rotorcraft damage tolerance methodology. 

8-1/8-2 
9. REFERENCES.

1. DOT/FAA/AR-95/9, February 1996, “Feasibility Study of a Rotorcraft Health and Usage 

Monitoring System (HUMS): Usage and Structural Life Monitoring Evaluation” and 

2. DOT/FAA/AR-95/50, February 1996, “Feasibility Study of a Rotorcraft Health and 

Usage Monitoring; system (HUMS): Results of Operator’s Evaluation.” 

3. Amendment 29-l 9 Advisory Circular, Number 29 -2A, Paragraph 621 “29.1309 (through 

Amendment 29-19) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.” 

4. DOCUMENT NO. RTCWDO-178B, “Software Considerations in Airborne Systems and 

Equipment Certification.” 

5. BHTI Report No. 412-910-001, “Flight Loads Survey for the Model 412 Helicopter.” 

6. BHTI Report No. 412-099-353, “Hazard Analysis for the Installation of 

Teledyne/Stewart Hughes HUMS in the Bell Model 412EP.” 

7. BHTI Report No. 412-910-006, “Fatigue Substantiation of the Dynamic Components for 

the Model 412 Helicopter.” 

9-1/9-2
APPENDIX A—412 HUMS USAGE PARAMETER ACCURACIES 

A-1 PRESSURE ALTITUDE ACCURACY. 

The performance specifications were obtained from section 2.2 of the Shadin Installation Manual 

(Rev A, Report Number 4028D dated August 20, 1998) in order to identify the error contribution 

of the air data computer. The accuracy analysis consisted of the following. 

• Accuracy ±25 ft from -1000 ft to 5000 ft 

 ±30 ft from 5,000 ft to 11,000 ft 

 ±40 ft from 11,000 ft to 30,000 ft 

For comparison purposes, the requirement specified in Title 14 Code of Federal Regulations 

(CFR) Part 135, Appendix C is ±100 ft to ±700 ft. 

Note: Altitude is not derived from an existing aircraft signal source, and subsequently, the 

requirement in note 1 of 14 CFR Part 135, Appendix C does not apply. 

A-2 INDICATED AIRSPEED ACCURACY. 

The performance specifications were obtained from section 2.2 of the Shadin Installation Manual 

(Rev A, Report Number 4028D dated August 20, 1998) in order to identify the error contribution 

of the air data computer. The accuracy analysis consisted of the following. 

• Accuracy ±1.0 Kt. 

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±5% or 

10 Kt. 

Note: Altitude rate is not derived from an existing aircraft signal source, and subsequently, the 

requirement in note 1 of 14 CFR Part 135, Appendix C does not apply. 

A-3 MAGNETIC HEADING ACCURACY. 

The performance specifications were obtained from the Honeywell TARSYN-H System

Maintenance Manual in order to identify the error contribution of the Directional Gyro. The 

performance information for the HUMS processing unit (HPU) was obtained from the Teledyne 

System Requirements Specification. The accuracy analysis consisted of the following: 

From table 2-11 of the Honeywell Maintenance Manual 22-15-03, page 2-51 dated June 1, 2000: 

• Accuracy ±2.0° (slaved accuracy) 

The input of the HPU (reference channel #34 of SRS 2237400-1 page C-3) is identified as a 

synchro-type input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 of 

SRS 2237400-1, sheet 32. All errors within the HPU identified as a root sum squared (RSS) 

value of ±0.17% full scale. 

A-1
The synchro input scaling within the HPU (reference SRS 2237400-1, sheet 18) is scaled for 

359.912° (4095 counts) full scale. Subsequently, the full-scale HPU accuracy is 

100%

0.17% × 359.912° = ±0.612° 

The system accuracy is then obtained by the following: 

System Accuracy (RSS) = ( ) ( ) o 0.612 2.0 2.09 2 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±5°. 

Note: The contribution of the HPU accuracy is 29% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

A-4 PITCH ATTITUDE ACCURACY. 

The performance specifications were obtained from the Honeywell TARSYN-H System

Maintenance Manual in order to identify the error contribution of the Vertical Gyro. The 

performance information for the HUMS processing unit (HPU) was obtained from the Teledyne 

System Requirements Specification. The accuracy analysis consisted of the following. 

From Table 2-11 of the Honeywell Maintenance Manual 22-15-03, page 2-51, dated June 1, 

2000: 

• Accuracy ±0.25° (labeled verticality error) 

Note: For validation of the accuracy, Honeywell engineering specification CP 2593996 was 

reviewed and the test limits verified the ±0.25°. 

The input of the HPU (reference channel #33 of SRS 2237400-1 page C-3) is identified as a 

synchro-type input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 of 

SRS 2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.17% full 

scale. 

The synchro input scaling within the HPU (reference SRS 2237400-1, sheet 18) is scaled for 

359.912° (4095 counts) full scale. Subsequently, the full-scale HPU accuracy is 

100%

0.17% × 359.912° = ±0.612°

The system accuracy was then obtained by the following 

System Accuracy RSS = ( ) ( ) o 0.612 2.0 0.66 2 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±2°. 

A-2
Note: The contribution of the HPU accuracy is 92.7% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

A-5 ROLL ATTITUDE ACCURACY. 

The performance specifications were obtained from the Honeywell TARSYN-H System

Maintenance Manual in order to identify the error contribution of the Vertical Gyro. The 

performance information for the HPU was obtained from the Teledyne System Requirements 

Specification. The accuracy analysis consisted of the following. 

From Table 2-11 of the Honeywell Maintenance Manual 22-15-03, page 2-51, dated June 1, 

2000: 

• Accuracy ±0.25° (labeled verticality error) 

Note: For validation of the accuracy, Honeywell engineering specification CP 2593996 was 

reviewed and the test limits verified the ±0.25° accuracy. 

The input of the HPU (reference channel #35 of SRS 2237400-1 page C-3) is identified as a 

synchro-type input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 of 

SRS 2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.17% full 

scale. 

The synchro input scaling within the HPU (reference SRS 2237400-1, sheet 18) is scaled for 

359.912° (4095 counts) full scale. Subsequently, the full-scale HPU accuracy is 

100%

0.17% × 359.912° = ±0.612°

The system accuracy was then obtained by the following 

System Accuracy RSS = ( ) ( ) o 0.612 0.25 0.66 2 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±2°

Note: The contribution of the HPU accuracy is 92.7% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

A-6 NORMAL ACCELERATION ACCURACY (Nz). 

The performance specifications were obtained from Patriot CMM, 31-33-15, revision dated 

05/94, page 3 component maintenance manual, in order to identify the error contribution of the 

accelerometer. The performance information for the HPU was obtained from the Teledyne 

System Requirements Specification. The accuracy analysis consisted of the following. 

A-3
• Accuracy Review of the accuracy requirement shows the sensor tolerance ±0.75% 

full scale. 

The sensor range specified is -3 to +6 g or 9 g

The sensor portion of the system accuracy is then: 

 Sensor Accuracy = 

100

0.75 × 9 g (full range) = ±0.068 g 

The additional sensor errors due to axial null and thermal null shift were 

stated as follows: 

 The scaling is then

g

Vdc

9.0

5.0 = 0.56 Vdc/g

Axial null tolerance specified = ±0.025 Vdc or 

Vdc g

Vdc

0.56 /

0.025

 = ±0.045 g

Thermal null shift = ±0.01% fs per degree F from -65° to 160°F 225°F × 

0.01 = ±2.25% 

or 

100

2.25 × 9 g = ±0.203 g

Sensor RSS accuracy (max. datum) = ( ) 0.045 (0.203) 0.208 g

2 2 + = ±

The input of the HPU (reference channel #1 (Normal) of SRS 2237400-1 page C-2) is identified 

as a low level DC (LLDC)-type input where the accuracy of that input type is specified in 

paragraph 3.2.1.2.3 of SRS 2237400-1, sheet 32. 

All errors within the HPU identified as an RSS value of ±0.2% full scale, which translates to 

 100

0.2 × 9 g = ±0.018 g

The LLDC input scaling within the HPU (reference SRS 2237400-1, sheet 18) is scaled for 

5 Vdc (4095 counts) full scale or 9 g. 

The system accuracy was then obtained by the following: 

System Accuracy RSS = ( ) 0.068 ( ) 0.018 0.07 g

2 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±0.2 g. 

A-4
The maximum datum error system accuracy was then obtained by the following 

Max Datum RSS = ( ) 0.208 ( ) 0.018 0.209 g

2 2 + = ±

For comparison, the requirement specified in 14 CFR Part 135, Appendix C is ±0.3 g maximum 

datum error. 

Note: The contribution of the HPU accuracy is 25.7% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

A-7 MAIN ROTOR SPEED ACCURACY. 

The performance requirements for the tachometer generators are identified in MIL-G-26611. 

Review of this document revealed that the error contribution of the tachometer generator is 

primarily mechanical (i.e., gearbox, transmission pad, etc.) relative to speed accuracy and, 

subsequently, beyond the scope of this analysis. The performance information for the HPU was 

obtained from the Teledyne System Requirements Specification. The accuracy analysis 

consisted of the following. 

The input of the HPU (reference channel frequency #1 of SRS 2237400-1 page C-4) is identified 

as a slow tach-type input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 

of SRS 2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.25% 

full scale. 

The slow tach (period measurement) input scaling within the HPU (reference SRS 2237400-1, 

sheet 18) is scaled from 7 Hz (1 count) to 250 Hz (4095 counts) full scale. As identified above, 

the full-scale HPU accuracy is 

±0.25% of 120% or ±0.3% full scale 

Because the error contribution is strictly due to the HPU, the HPU accuracy becomes the total 

system accuracy. 

For comparison, the requirement specified in 14 CFR Part 135, Appendix C is ±5%. 

Note: The contribution of the HPU accuracy is 100% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

A-8 ENGINE/MAST TORQUE ACCURACY—ENGINES 1&2& MAST. 

The following paragraphs identify the accuracy, range, sampling interval, and resolution for 

engine torque. 

1. Engine Torque—Bell 412 serial numbers (S/N) 33001 through 33213 and 36001 through 

36019 

• Engine Torque Accuracy 

A-5
Note: This sensor is installed in the above referenced S/N aircraft except for those aircraft 

modified with the improved hover performance kit (Bell P/N 412-570-001-103). 

The performance specifications were obtained from Courter Inc. (Bendix) 172-1842-7 

corresponding to Bell Sensor P/N specification 412-075-205 in order to identify the error 

contribution of the torque pressure transducer. The performance information for the HPU was 

obtained from the Teledyne System Requirements Specification. The accuracy analysis 

consisted of the following 

From Courter Inc. (Bendix) drawing 172-1842-7, Revision N, dated 10/21/82, Table I: 

• Accuracy 

Note: Review of the accuracy presented in Table I shows a bell shape curve with accuracies 

ranging from ±2.5° @ 0 psi, ±0.7° @ 35psi to ±2.2° @ 65 psi. The accuracies were presented as 

percent of point rather than percent of full scale. Given that the range generally will be between 

20 psi and higher, and to be conservative, the tolerance at that point (20 psi) was used (±1.5°). 

Therefore: 

The sensor accuracy is then 69.6

1.5 × 100 = ±2.16% 

The input of the HPU (reference channels #42 and #43 of SRS 2237400-1 page C-3) is identified 

as a synchro-type input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 

of SRS 2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.17% 

full scale. 

The synchro input scaling within the HPU (reference SRS 2237400-1, sheet 18) is scaled for 

359.912° (4095 counts) full scale. 

The system accuracy was then obtained by the following 

System Accuracy RSS = ( ) 2.16 ( ) 0.17 2.167% 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±5%. 

Note: The contribution of the HPU accuracy is 7.8% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

A-9 OUTSIDE AIR TEMPERATURE ACCURACY. 

The performance specifications were obtained from Shadin in order to identify the error 

contribution of the air data computer. The accuracy analysis consisted of the following. 

A-6
From section 2.2 of the Shadin Installation Manual (Rev A, Report Number 4028D dated August 

20, 1998) and supported by additional data from Shadin: 

• Accuracy ±2°C 

A-10 ALTITUDE RATE ACCURACY. 

The performance specifications were obtained from Shadin in order to identify the error 

contribution of the Air data computer. The accuracy analysis consisted of the following. 

From section 2.2 of the Shadin Installation Manual (Rev A, Report Number 4028D dated August 

20, 1998) and supported by additional data from Shadin: 

• Accuracy ±40 ft/min 

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is 

±250 ft/min below 12,000 ft altitude and ±10% above 12,000 ft altitude. 

Note: Altitude rate is not derived from an existing aircraft signal source, and subsequently, the 

requirement in note 1 of 14 CFR Part 135, Appendix C does not apply. 

A-11 COLLECTIVE POSITION ACCURACY. 

The performance specifications were obtained from the Bell Sensor drawing P/N 412-074-101 in 

order to identify the error contribution of the control motion transducer. The performance 

information for the HPU was obtained from the Teledyne System Requirements Specification. 

The accuracy analysis consisted of the following. 

From Bell Specification 412-074-101, Rev E, dated 2/11/93: 

• Accuracy Review of the performance data in Note 2, Item A1, shows the resistance 

tolerance over the entire environmental range is ±15%. The linearity 

tolerance, indicated in A-2 is ±0.5% over the entire operating range. 

The nonrigged sensor accuracy is then 

Sensor RSS accuracy (nonrigged) = ( ) 15.0 (0.5) 15.01% 2 2 + = ±

Note: The accuracy above does not include the mechanical misalignment due to mechanical 

tolerance in the sensor package. 

The input of the HPU (reference channel #6 of SRS 2237400-1 page C-2) is identified as a 

potentiometer (POT)-type input where the accuracy of that input type is specified in paragraph 

3.2.1.2.3 of SRS 2237400-1, sheet 32. All errors within the HPU identified as an RSS value of 

±0.2% full scale. 

A-7
The POT input scaling within the HPU (reference SRS 2237400-1, sheet 19) is scaled for 5 Vdc 

(4095 counts) full scale. 

The system accuracy was then obtained by the following. 

System Accuracy RSS = ( ) 15.01 (0.2) 15.01% 2 2 + = ±

For comparison, the requirement specified in 14 CFR Part 135, Appendix C is ±3%. 

Note: The contribution of the HPU accuracy is 1.3% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

Installed Accuracy: The Collective position sensor is rigged per procedure to be within ±3% of 

Travel. 

A-12 F/A CYCLIC POSITION (LONGITUDINAL) ACCURACY. 

The performance specifications were obtained from the Bell Sensor drawing P/N 412-074-101 in 

order to identify the error contribution of the control motion transducer. The performance 

information for the HPU was obtained from the Teledyne System Requirements Specification. 

The accuracy analysis consisted of the following. 

From Bell Specification 412-074-101, Rev E, dated 2/11/93: 

Accuracy Review of the performance data in Note 2, Item A1, shows the resistance tolerance 

over the entire environmental range is ±15%. The linearity tolerance, indicated in 

A-2 is ±0.5% over the entire operating range. 

The nonrigged sensor accuracy is then 

Sensor RSS accuracy (nonrigged) = ( ) 15.01 (0.2) 15.01% 2 2 + = ±

Note: The accuracy above does not include the mechanical misalignment due to mechanical 

tolerance in the sensor package. 

The input of the HPU (reference channel #6 of SRS 2237400-1 page C-2) is identified as a POTtype input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 of SRS 

2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.2% full scale. 

The POT input scaling within the HPU (reference SRS 2237400-1, sheet 19) is scaled for 5 Vdc 

(4095 counts) full scale. 

The system accuracy was then obtained by the following 

System Accuracy RSS = ( ) 15.01 (0.2) 15.01% 2 2 + = ±

A-8
For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±3%. 

Note: The contribution of the HPU accuracy is 1.3% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

Installed Accuracy: The Cyclic position sensor is rigged per procedure to be within ±3% of 

Travel. 

A-13 CYCLIC POSITION (LATERAL) ACCURACY. 

The performance specifications were obtained from the Bell Sensor drawing P/N 412-074-101 in 

order to identify the error contribution of the control motion transducer. The performance 

information for the HPU was obtained from the Teledyne System Requirements Specification. 

The accuracy analysis consisted of the following. 

From Bell Specification 412-074-101, Rev E, dated 2/11/93: 

• Accuracy Review of the performance data in Note 2, Item A1, shows the resistance 

tolerance over the entire environmental range is ±15%. The linearity 

tolerance, indicated in A-2 is ±0.5% over the entire operating range. 

The nonrigged sensor accuracy is then 

Sensor RSS accuracy (nonrigged) = ( ) 15.01 (0.5) 15.01% 2 2 + = ±

Note: The accuracy above does not include the mechanical misalignment due to mechanical 

tolerance in the sensor package. 

The input of the HPU (reference channel #6 of SRS 2237400-1 page C-2) is identified as a POTtype input were the accuracy of that input type is specified in paragraph 3.2.1.2.3 of SRS 

2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.2% full scale. 

The POT input scaling within the HPU (reference SRS 2237400-1, sheet 19) is scaled for 5 Vdc 

(4095 counts) full scale. 

The system accuracy was then obtained by the following 

System Accuracy RSS = ( ) 15.01 (0.2) 15.01% 2 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±3%. 

Note: The contribution of the HPU accuracy is 1.3% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

Installed Accuracy: The Cyclic position sensor is rigged per procedure to be within ±3% of 

Travel. 

A-9
A-14 PEDAL POSITION ACCURACY. 

The performance specifications were obtained from the Bell Sensor drawing P/N 412-074-101 in 

order to identify the error contribution of the control motion transducer. The performance 

information for the HPU was obtained from the Teledyne System Requirements Specification. 

The accuracy analysis consisted of the following. 

From Bell Specification 412-074-101, Rev E, dated 2/11/93: 

Accuracy Review of the performance data in Note 2, Item A1, shows the resistance tolerance 

over the entire environmental range is ±15%. The linearity tolerance, indicated in 

A-2 is ±0.5% over the entire operating range. 

The nonrigged sensor accuracy is then: 

Sensor RSS accuracy (nonrigged) = ( ) 15.01 (0.5) 15.01% 2 2 + = ±

Note: The accuracy above does not include the mechanical misalignment due to mechanical 

tolerance in the sensor package. 

The input of the HPU (reference channel #6 of SRS 2237400-1 page C-2) is identified as a POTtype input where the accuracy of that input type is specified in paragraph 3.2.1.2.3 of SRS 

2237400-1, sheet 32. All errors within the HPU identified as an RSS value of ±0.2% full scale. 

The POT input scaling within the HPU (reference SRS 2237400-1, sheet 19) is scaled for 5 Vdc 

(4095 counts) full scale. 

The system accuracy was then obtained by the following 

System Accuracy RSS = ( ) 15.01 (0.2) 15.01% 2 2 + = ±

For comparison purposes, the requirement specified in 14 CFR Part 135, Appendix C is ±3%. 

Note: The contribution of the HPU accuracy is 1.3% in reference to note 1 in 14 CFR Part 135, 

Appendix C. 

Installed Accuracy: The Pedal position sensor is rigged per procedure to be within ±3% of 

Travel. 

A-10
A-15 TIME AND DATE ACCURACY. 

Time is provided by a time clock circuit internal to the Teledyne HPU. Subsequently the time 

accuracy is solely associated to the internal HPU circuit. The accuracy of the time circuit was 

provided by Teledyne and is as follows 

Installed system accuracy = ±0.125% per hour or ±4.5 sec 

For comparison, the requirement specified in 14 CFR Part 135, Appendix C is ±0.125% per 

hour. 

Note: The time signal is not derived from an existing aircraft signal source, and subsequently, 

the requirement in note 1 of 14 CFR Part 135, Appendix C does not apply. 

Time/Date Setup: Note the time and date are set by the operator using a HPU front panel 

function. 

A-16 GROSS WEIGHT. 

Gross weight is a pilot-entered value. It is set to the nearest 50 pounds. 

A-17 ON GROUND DISCRETE. 

Validates each flight leg. Accuracy is not applicable. 

A-18 AIRCRAFT FLIGHT HOURS.

Cumulative aircraft flight hours are operator entered and tracked by the HUMS system for 

system integrity checking. Usage data is not used if the current value does not match the 

aircraft’s logbook. 

A-11/A-12
APPENDIX B—ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh h % Error nstart nend

3 308

-55 0 8000 0.870103 -9821.29 0.104057

-50 0 8000 0.880018 -8997.33 0.109118

-45 0 8000 0.889822 -8195.86 0.113382

-40 0 8000 0.899519 -7415.79 0.116892

-35 0 8000 0.909112 -6656.11 0.119728

-30 0 8000 0.918606 -5915.88 0.121944

-25 0 8000 0.928002 -5194.21 0.123573

-20 0 8000 0.937304 -4490.28 0.12467

-15 0 8000 0.946515 -3803.31 0.125284

-10 0 8000 0.955637 -3132.58 0.125436

-5 0 8000 0.964673 -2477.41 0.125182

0 0 8000 0.973625 -1837.15 0.124531

5 0 8000 0.982495 -1211.2 0.12352

10 0 8000 0.991286 -598.998 0.123478

15 0 8000 1 0 0.129877

20 0 8000 1.008638 586.2985 0.136235

25 0 8000 1.017204 1160.374 0.142558

-55 0 10000 0.870103 -9821.29 0.201197

-50 0 10000 0.880018 -8997.33 0.203273

-45 0 10000 0.889822 -8195.86 0.205106

-40 0 10000 0.899519 -7415.79 0.206717

-35 0 10000 0.909112 -6656.11 0.208116

-30 0 10000 0.918606 -5915.88 0.209325

-25 0 10000 0.928002 -5194.21 0.210346

-20 0 10000 0.937304 -4490.28 0.211202

-15 0 10000 0.946515 -3803.31 0.211892

-10 0 10000 0.955637 -3132.58

dV

129.6284

129.4333

129.2224

128.997

128.7584

128.5078

128.2462

127.9748

127.6943

127.4056

127.1094

126.8065

126.4974

126.1827

125.863

125.5387

125.2103

130.248

129.7735

129.3054

128.8436

128.3881

127.9387

127.4952

127.0577

126.6259

126.1998 0.212439

-5 0 10000 0.964673 -2477.41 125.7792 0.212858

0 0 10000 0.973625 -1837.15 125.364 0.213137

5 0 10000 0.982495 124.9542 0.213298

10 0 10000 0.991286 -598.998 124.5495 0.213348

15 0 10000 1 0 124.15 0.213285

20 0 10000 1.008638 586.2985 123.7555 0.21312

25 0 10000 1.017204 1160.374 123.3658 0.21286

-55 0 11900 0.870103 -9821.29 125.4123 0.224303

-50 0 11900 0.880018 -8997.33 125.4101 0.177622

-45 0 11900 0.889822 -8195.86 125.3439 0.174769

-40 0 11900 0.899519 -7415.79 125.2187 0.174122

-35 0 11900 0.909112 -6656.11 125.0392 0.175543

-30 0 11900 0.918606 -5915.88 124.8096 0.178916

-25 0 11900 0.928002 -5194.21 124.5339 0.184134

-1211.2

B-1
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

-20 0 11900 0.937304 -4490.28 124.2155 0.19127

-15 0 11900 0.946515 -3803.31 123.8578 0.201549

-10 0 11900 0.955637 -3132.58 123.4637 0.21342

-5 0 11900 0.964673 -2477.41 123.036 0.2268

0 0 11900 0.973625 -1837.15 122.5771 0.24164

5 0 11900 0.982495 -1211.2 122.0896 0.257872

10 0 11900 0.991286 -598.998 121.5754 0.275448

15 0 11900 1 0 121.0366 0.294336

20 0 11900 1.008638 586.2985 120.475 0.314472

25 0 11900 1.017204 1160.374 119.8923 0.335836

-55 3000 8000 0.919125 -5875.72 128.4937 0.131695

-50 3000 8000 0.929598 -5072.71 128.2005 0.133877

-45 3000 8000 0.939955 -4291.62 127.8953 0.13536

-40 3000 8000 0.950198 -3531.39 127.579 0.136221

-35 3000 8000 0.960332 -2791.03 127.2529 0.136511

-30 3000 8000 0.970361 -2069.62 126.918 0.136264

-25 3000 8000 0.980286 -1366.3 126.5751 0.135536

-20 3000 8000 0.990113 -680.267 126.2252 0.134975

-15 3000 8000 0.999842 -10.7626 125.8688 0.142364

-10 3000 8000 1.009478 642.9112 125.5068 0.14968

-5 3000 8000 1.019023 1281.426 125.1397 0.156947

0 3000 8000 1.028479 1905.406 124.7681 0.164147

5 3000 8000 1.03785 2515.437 124.3926 0.171304

10 3000 8000 1.047136 3112.074 124.0136 0.17841

15 3000 8000 1.05634 3695.84 123.6315 0.18549

20 3000 8000 1.065466 4267.229 123.2467 0.192532

25 3000 8000 1.074513 4826.709 122.8597 0.199553

-55 3000 10000 0.919125 -5875.72 127.9141 0.226083

-50 3000 10000 0.929598 -5072.71 127.4201 0.227014

-45 3000 10000 0.939955 -4291.62 126.9333 0.227752

-40 3000 10000 0.950198 -3531.39 126.4537 0.228296

-35 3000 10000 0.960332 -2791.03 125.981 0.228668

-30 3000 10000 0.970361 -2069.62 125.5152 0.228872

-25 3000 10000 0.980286 -1366.3 125.0561 0.228932

-20 3000 10000 0.990113 -680.267 124.6035 0.228845

-15 3000 10000 0.999842 -10.7626 124.1572 0.228629

-10 3000 10000 1.009478 642.9112 123.7172 0.228295

-5 3000 10000 1.019023 1281.426 123.2832 0.227842

0 3000 10000 1.028479 1905.406 122.8552 0.227282

5 3000 10000 1.03785 2515.437 122.433 0.226627

10 3000 10000 1.047136 3112.074 122.0165 0.225868

15 3000 10000 1.05634 3695.84 121.6055 0.225032

B-2
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

20 3000 10000 1.065466 4267.229 121.1999 0.22411

25 3000 10000 1.074513 4826.709 120.7996 0.22311

-55 3000 11900 0.919125 -5875.72 124.7956 0.188724

-50 3000 11900 0.929598 -5072.71 124.4824 0.196596

-45 3000 11900 0.939955 -4291.62 124.1169 0.208037

-40 3000 11900 0.950198 -3531.39 123.7034 0.221857

-35 3000 11900 0.960332 -2791.03 123.246 0.237488

-30 3000 11900 0.970361 -2069.62 122.7484 0.254859

-25 3000 11900 0.980286 -1366.3 122.214 0.273903

-20 3000 11900 0.990113 -680.267 121.6458 0.294549

-15 3000 11900 0.999842 -10.7626 121.0465 0.316744

-10 3000 11900 1.009478 642.9112 120.4189 0.34043

-5 3000 11900 1.019023 1281.426 119.7653 0.365578

0 3000 11900 1.028479 1905.406 119.0877 0.392125

5 3000 11900 1.03785 2515.437 118.3882 0.420052

10 3000 11900 1.047136 3112.074 117.6686 0.449321

15 3000 11900 1.05634 3695.84 116.9307 0.479899

20 3000 11900 1.065466 4267.229 116.1759 0.511768

25 3000 11900 1.074513 4826.709 115.4056 0.544893

-55 6000 8000 0.972043 -1949.64 126.8607 0.149267

-50 6000 8000 0.983119 -1167.48 126.4753 0.14889

-45 6000 8000 0.994072 -406.669 126.0813 0.152169

-40 6000 8000 1.004905 333.827 125.6797 0.160692

-35 6000 8000 1.015623 1054.966 125.2714 0.169103

-30 6000 8000 1.026229 1757.648 124.8573 0.17743

-25 6000 8000 1.036726 2442.708 124.438 0.185661

-20 6000 8000 1.047118 3110.93 124.0143 0.193826

-15 6000 8000 1.057408 3763.049 123.5868 0.201923

-10 6000 8000 1.067599 4399.754 123.156 0.20997

-5 6000 8000 1.077693 5021.693 122.7224 0.217961

0 6000 8000 1.087694 5629.473 122.2865 0.225912

5 6000 8000 1.097603 6223.666 121.8488 0.233831

10 6000 8000 1.107424 6804.813 121.4094 0.241715

15 6000 8000 1.117159 7373.423 120.9689 0.24957

20 6000 8000 1.126809 7929.977 120.5276 0.257403

25 6000 8000 1.136378 8474.932 120.0856 0.265225

-55 6000 10000 0.972043 -1949.64 125.4373 0.246854

-50 6000 10000 0.983119 -1167.48 124.9254 0.246674

-45 6000 10000 0.994072 -406.669 124.4216 0.246318

-40 6000 10000 1.004905 333.827 123.9258 0.24582

-35 6000 10000 1.015623 1054.966 123.4376 0.245173

-30 6000 10000 1.026229 1757.648 122.9569 0.244394

B-3
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

-25 6000 10000 1.036726 2442.708 122.4836 0.243488

-20 6000 10000 1.047118 3110.93 122.0173 0.242474

-15 6000 10000 1.057408 3763.049 121.558 0.241351

-10 6000 10000 1.067599 4399.754 121.1054 0.240136

-5 6000 10000 1.077693 5021.693 120.6594 0.238823

0 6000 10000 1.087694 5629.473 120.2198 0.237424

5 6000 10000 1.097603 6223.666 119.7865 0.235952

10 6000 10000 1.107424 6804.813 119.3592 0.2344

15 6000 10000 1.117159 7373.423 118.938 0.232786

20 6000 10000 1.126809 7929.977 118.5225 0.231098

25 6000 10000 1.136378 8474.932 118.1127 0.229348

-55 6000 11900 0.972043 -1949.64 122.6607 0.27971

-50 6000 11900 0.983119 -1167.48 122.0541 0.303172

-45 6000 11900 0.994072 -406.669 121.4062 0.328493

-40 6000 11900 1.004905 333.827 120.7209 0.355625

-35 6000 11900 1.015623 1054.966 120.0017 0.384497

-30 6000 11900 1.026229 1757.648 119.2516 0.415059

-25 6000 11900 1.036726 2442.708 118.4735 0.447273

-20 6000 11900 1.047118 3110.93 117.6701 0.481105

-15 6000 11900 1.057408 3763.049 116.8436 0.516503

-10 6000 11900 1.067599 4399.754 115.9962 0.553452

-5 6000 11900 1.077693 5021.693 115.13 0.591923

0 6000 11900 1.087694 5629.473 114.2467 0.631893

5 6000 11900 1.097603 6223.666 113.3479 0.673344

10 6000 11900 1.107424 6804.813 112.4353 0.716259

15 6000 11900 1.117159 7373.423 111.5102 0.760625

20 6000 11900 1.126809 7929.977 110.5739 0.806437

25 6000 11900 1.136378 8474.932 109.6275 0.853693

-55 9000 8000 1.029261 1956.627 124.7371 0.196996

-50 9000 8000 1.040989 2718.044 124.2652 0.206475

-45 9000 8000 1.052587 3458.681 123.7881 0.215844

-40 9000 8000 1.064058 4179.538 123.3065 0.225102

-35 9000 8000 1.075406 4881.555 122.8212 0.234284

-30 9000 8000 1.086636 5565.6 122.3329 0.243389

-25 9000 8000 1.097751 6232.492 121.8422 0.252428

-20 9000 8000 1.108755 6882.994 121.3495 0.261406

-15 9000 8000 1.119651 7517.82 120.8554 0.270341

-10 9000 8000 1.130442 8137.641 120.3603 0.279237

-5 9000 8000 1.14113 8743.085 119.8646 0.288093

0 9000 8000 1.151719 9334.747 119.3686 0.296921

5 9000 8000 1.162212 9913.184 118.8727 0.305727

10 9000 8000 1.172611 10478.92 118.3772 0.314883

B-4
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

15 9000 8000 1.182919 11032.45 117.8822 0.324547

20 9000 8000 1.193138 11574.25 117.388 0.33422

25 9000 8000 1.203269 12104.75 116.8948 0.343912

-55 9000 10000 1.029261 1956.627 122.8199 0.26342

-50 9000 10000 1.040989 2718.044 122.292 0.26213

-45 9000 10000 1.052587 3458.681 121.7729 0.260704

-40 9000 10000 1.064058 4179.538 121.2624 0.259159

-35 9000 10000 1.075406 4881.555 120.7602 0.257495

-30 9000 10000 1.086636 5565.6 120.2662 0.255736

-25 9000 10000 1.097751 6232.492 119.78 0.253882

-20 9000 10000 1.108755 6882.994 119.3015 0.251933

-15 9000 10000 1.119651 7517.82 118.8305 0.249908

-10 9000 10000 1.130442 8137.641 118.3667 0.247806

-5 9000 10000 1.14113 8743.085 117.9099 0.245634

0 9000 10000 1.151719 9334.747 117.46 0.243399

5 9000 10000 1.162212 9913.184 117.0168 0.2411

10 9000 10000 1.172611 10478.92 116.5801 0.238751

15 9000 10000 1.182919 11032.45 116.1497 0.236351

20 9000 10000 1.193138 11574.25 115.7254 0.233901

25 9000 10000 1.203269 12104.75 115.3072 0.231408

-55 9000 11900 1.029261 1956.627 119.0304 0.458179

-50 9000 11900 1.040989 2718.044 118.1478 0.496498

-45 9000 11900 1.052587 3458.681 117.2345 0.536725

-40 9000 11900 1.064058 4179.538 116.2938 0.578822

-35 9000 11900 1.075406 4881.555 115.3285 0.622763

-30 9000 11900 1.086636 5565.6 114.3412 0.668529

-25 9000 11900 1.097751 6232.492 113.3343 0.716098

-20 9000 11900 1.108755 6882.994 112.31 0.765446

-15 9000 11900 1.119651 7517.82 111.2702 0.816571

-10 9000 11900 1.130442 8137.641 110.2167 0.86946

-5 9000 11900 1.14113 8743.085 109.1511 0.924115

0 9000 11900 1.151719 9334.747 108.0749 0.98053

5 9000 11900 1.162212 9913.184 106.9895 1.038712

10 9000 11900 1.172611 10478.92 105.896 1.098674

15 9000 11900 1.182919 11032.45 104.7956 1.160408

20 9000 11900 1.193138 11574.25 103.6893 1.223934

25 9000 11900 1.203269 12104.75 102.5781 1.289279

-55 12000 8000 1.091235 5842.753 122.1307 0.268161

-50 12000 8000 1.10367 6583.535 121.578 0.278479

-45 12000 8000 1.115965 7304.099 121.0232 0.28872

-40 12000 8000 1.128127 8005.419 120.467 0.298882

-35 12000 8000 1.140159 8688.409 119.9099 0.308981

B-5
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

-30 12000 8000 1.152065 9353.918 119.3524 0.319015

-25 12000 8000 1.163849 10002.74 118.795 0.329015

-20 12000 8000 1.175516 10635.61 118.2381 0.338979

-15 12000 8000 1.187067 11253.23 117.6819 0.349716

-10 12000 8000 1.198508 11856.25 117.127 0.360668

-5 12000 8000 1.20984 12445.29 116.5734 0.371628

0 12000 8000 1.221067 13020.91 116.0216 0.382596

5 12000 8000 1.232192 13583.67 115.4716 0.393575

10 12000 8000 1.243217 14134.08 114.9237 0.404573

15 12000 8000 1.254145 14672.61 114.3781 0.415588

20 12000 8000 1.264979 15199.72 113.8349 0.426626

25 12000 8000 1.275721 15715.84 113.2942 0.437679

-55 12000 10000 1.091235 5842.753 120.0647 0.275641

-50 12000 10000 1.10367 6583.535 119.5223 0.273241

-45 12000 10000 1.115965 7304.099 118.9895 0.270739

-40 12000 10000 1.128127 8005.419 118.4659 0.268155

-35 12000 10000 1.140159 8688.409 117.9513 0.265483

-30 12000 10000 1.152065 9353.918 117.4454 0.262736

-25 12000 10000 1.163849 10002.74 116.9479 0.259919

-20 12000 10000 1.175516 10635.61 116.4586 0.257042

-15 12000 10000 1.187067 11253.23 115.9772 0.254102

-10 12000 10000 1.198508 11856.25 115.5035 0.251115

-5 12000 10000 1.20984 12445.29 115.0372 0.248074

0 12000 10000 1.221067 13020.91 114.5783 0.244993

5 12000 10000 1.232192 13583.67 114.1264 0.241871

10 12000 10000 1.243217 14134.08 113.6813 0.238704

15 12000 10000 1.254145 14672.61 113.2429 0.235512

20 12000 10000 1.264979 15199.72 112.811 0.232282

25 12000 10000 1.275721 15715.84 112.3854 0.229027

-55 12000 11900 1.091235 5842.753 113.9281 0.735382

-50 12000 11900 1.10367 6583.535 112.7867 0.791251

-45 12000 11900 1.115965 7304.099 111.6247 0.849307

-40 12000 11900 1.128127 8005.419 110.4446 0.909514

-35 12000 11900 1.140159 8688.409 109.2488 0.971894

-30 12000 11900 1.152065 9353.918 108.0395 1.036449

-25 12000 11900 1.163849 10002.74 106.8185 1.103185

-20 12000 11900 1.175516 10635.61 105.5876 1.172131

-15 12000 11900 1.187067 11253.23 104.3483 1.243286

-10 12000 11900 1.198508 11856.25 103.1021 1.316677

-5 12000 11900 1.20984 12445.29 101.8502 1.392345

0 12000 11900 1.221067 13020.91 100.5938 1.470304

5 12000 11900 1.232192 13583.67 99.33396 1.550603

B-6
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

10 12000 11900 1.243217 14134.08 98.07162 1.633272

15 12000 11900 1.254145 14672.61 96.80765 1.718346

20 12000 11900 1.264979 15199.72 95.54282 1.805895

25 12000 11900 1.275721 15715.84 94.27785 1.895957

-55 15000 8000 1.158481 9708.388 119.0496 0.349537

-50 15000 8000 1.171682 10428.64 118.4216 0.3609

-45 15000 8000 1.184735 11129.24 117.7946 0.372193

-40 15000 8000 1.197646 11811.13 117.1689 0.384293

-35 15000 8000 1.210419 12475.19 116.545 0.396719

-30 15000 8000 1.223059 13122.26 115.9233 0.409147

-25 15000 8000 1.23557 13753.1 115.304 0.421581

-20 15000 8000 1.247955 14368.44 114.6874 0.434025

-15 15000 8000 1.260219 14968.94 114.0738 0.446479

-10 15000 8000 1.272364 15555.26 113.4634 0.458947

-5 15000 8000 1.284395 16127.97 112.8562 0.471448

0 15000 8000 1.296313 16687.65 112.2526 0.483954

5 15000 8000 1.308124 17234.81 111.6526 0.496504

10 15000 8000 1.319828 17769.97 111.0563 0.509069

15 15000 8000 1.33143 18293.57 110.4638 0.521662

20 15000 8000 1.342931 18806.08 109.8753 0.534289

25 15000 8000 1.354335 19307.9 109.2907 0.546955

-55 15000 10000 1.158481 9708.388 117.1741 0.283327

-50 15000 10000 1.171682 10428.64 116.619 0.279814

-45 15000 10000 1.184735 11129.24 116.0741 0.276238

-40 15000 10000 1.197646 11811.13 115.539 0.272598

-35 15000 10000 1.210419 12475.19 115.0135 0.268901

-30 15000 10000 1.223059 13122.26 114.4971 0.265163

-25 15000 10000 1.23557 13753.1 113.9897 0.261384

-20 15000 10000 1.247955 14368.44 113.4909 0.257558

-15 15000 10000 1.260219 14968.94 113.0005 0.253707

-10 15000 10000 1.272364 15555.26 112.5182 0.249811

-5 15000 10000 1.284395 16127.97 112.0437 0.245884

0 15000 10000 1.296313 16687.65 111.5769 0.241942

5 15000 10000 1.308124 17234.81 111.1175 0.237978

10 15000 10000 1.319828 17769.97 110.6653 0.233979

15 15000 10000 1.33143 18293.57 110.22 0.229982

20 15000 10000 1.342931 18806.08 109.7816 0.22596

25 15000 10000 1.354335 19307.9 109.3497 0.221926

-55 15000 11900 1.158481 9708.388 107.3776 1.135184

-50 15000 11900 1.171682 10428.64 105.9944 1.213346

-45 15000 11900 1.184735 11129.24 104.6001 1.294261

-40 15000 11900 1.197646 11811.13 103.1966 1.377965

B-7
ERROR ANALYSIS OF Vh FOR THE BELL HELICOPTER MODEL 412 (Continued) 

OAT (°C) Palt GWT 1/sqr(sig) DenAlt Vh dVh % Error nstart nend

-35 15000 11900 1.210419 12475.19 101.7857 1.46451

-30 15000 11900 1.223059 13122.26 100.3692 1.55394

-25 15000 11900 1.23557 13753.1 98.94855 1.6463

-20 15000 11900 1.247955 14368.44 97.52507 1.741683

-15 15000 11900 1.260219 14968.94 96.09995 1.840138

-10 15000 11900 1.272364 15555.26 94.67425 1.94173

-5 15000 11900 1.284395 16127.97 93.24894 2.046542

0 15000 11900 1.296313 16687.65 91.82489 2.154695

5 15000 11900 1.308124 17234.81 90.40285 2.266247

10 15000 11900 1.319828 17769.97 88.98351 2.381284

15 15000 11900 1.33143 18293.57 87.56752 2.499972

20 15000 11900 1.342931 18806.08 86.1554 2.622373

25 15000 11900 1.354335 19307.9 84.74766 2.748657

B-8