Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 49155

FAA finds that good cause exists to 

make these special conditions effective 

upon issuance.

List of Subjects in 14 CFR Part 25

Aircraft, Aviation safety, Reporting 

and recordkeeping requirements.

The authority citation for these 

special conditions is as follows:

Authority: 49 U.S.C. 106(g), 40113, 44701, 

44702, 44704. 

The Special Conditions 

Accordingly, pursuant to the 

authority delegated to me by the 

Administrator, the following special 

conditions are issued as part of the type 

certification basis for the Boeing Model 

777 airplane. 

In addition to the airworthiness 

standards of §§ 25.562 and 25.785, the 

minimum acceptable standards for 

dynamic certification of Boeing Model 

777 single-occupant side-facing seats are 

as follows:

Additional Injury Criteria 

(a) Existing Criteria: All injury 

protection criteria of § 25.562(c)(1) 

through (c)(6) apply to the occupant of 

a side-facing seat. Head Injury Criterion 

(HIC) assessments are only required for 

head contact with the seat and/or 

adjacent structures. 

(b) Body-to-Wall/Furnishing Contact:

Under the load condition defined in 

§ 25.562(b)(2), the seat must be installed 

immediately aft of a structure such as an 

interior wall or furnishing that will 

support the pelvis, upper arm, chest, 

and head of an occupant seated next to 

the structure. A conservative 

representation of the structure and its 

stiffness must be included in the tests. 

It is recommended, but not required, 

that the contact surface of this structure 

be covered with at least two inches of 

energy absorbing protective padding 

(foam or equivalent), such as Ensolite. 

(c) Thoracic Trauma: Under the load 

condition defined in § 25.562(b)(2), 

Thoracic Trauma Index (TTI) injury 

criterion must be substantiated by 

dynamic test or by rational analysis 

based on previous test(s) of a similar 

seat installation. Testing must be 

conducted with a Side Impact Dummy 

(SID), as defined by Title 49 Code of 

Federal Regulations (CFR) part 572, 

Subpart F, or its equivalent. The TTI 

must be less than 85, as defined in 49 

CFR part 572, Subpart F. The SID TTI 

data must be processed as defined in 

Federal Motor Vehicle Safety Standard 

(FMVSS) part 571.214, section S6.13.5. 

(d) Pelvis: Under the load condition 

defined in § 25.562(b)(2), pelvic lateral 

acceleration must be shown by dynamic 

test or by rational analysis based on 

previous test(s) of a similar seat 

installation to not exceed 130g. Pelvic 

acceleration data must be processed as 

defined in FMVSS part 571.214, section 

S6.13.5. 

(e) Shoulder Strap Loads: Where 

upper torso straps (shoulder straps) are 

used for occupants, tension loads in 

individual straps must not exceed 1,750 

pounds. If dual straps are used for 

restraining the upper torso, the total 

strap tension loads must not exceed 

2,000 pounds. 

(f) Neck Injury Criteria: The seating 

system must protect the occupant from 

experiencing serious neck injury. 

Inflatable Lapbelt Conditions 

(a) If inflatable lapbelts are used as the 

means of occupant restraint on single 

place side-facing seats, the requirements 

of existing Special Conditions 25–04–

03–SC (1–14), ‘‘Boeing Model 777 Series 

Airplanes; Seats with Inflatable 

Lapbelts’’ are incorporated by reference 

except for special conditions 1 and 3, 

which are replaced by (b) and (c) below. 

(b) Seats With Inflatable Lapbelts. It 

must be shown that the inflatable 

lapbelt will deploy and provide 

protection under crash conditions 

where it is necessary to prevent serious 

head, neck, thoracic, and pelvic lateral 

acceleration injury from body-to-wall/

furnishing contact. The means of 

protection must take into consideration 

a range of stature from two-year-old 

child to ninety-fifth percentile male. 

The inflatable lapbelt must provide a 

consistent approach to energy 

absorption throughout the range. In 

addition, the following situations must 

be considered: 

1. The seat occupant is holding an 

infant. 

2. The seat occupant is a child in a 

child restraint device. 

3. The seat occupant is a child not 

using a child restraint device. 

4. The seat occupant is a pregnant 

woman. 

(c) The design must prevent the 

inflatable lapbelt from being either 

incorrectly buckled or incorrectly 

installed such that the inflatable lapbelt 

would not properly deploy. 

Alternatively, it must be shown that 

such deployment is not hazardous to the 

occupant, and will provide the required 

injury protection.

Note: The existing means of controlling 

HIC, TTI and pelvic lateral acceleration result 

in a progressive reduction of injury severity 

for impact conditions less than the maximum 

specified by the requirements. However, 

airbag technology involves a step change in 

protection for impacts below and above that 

at which the airbag deploys. This could 

result in one or more of the injury criteria 

being higher at an intermediate impact 

condition than that resulting from the 

maximum. The step change in injury 

protection is acceptable, provided that the 

injury criteria values for any intermediate 

impact (whether or not the inflatable lapbelt 

delays) do not exceed the maximum allowed 

by the requirements.

Additional Test Requirements 

(a) One longitudinal test with the SID 

Anthropomorphic Test Dummy (ATD), 

undeformed floor, no yaw, and with all 

lateral structural supports (armrests/

walls). 

Pass/fail injury assessments: The TTI 

and pelvic acceleration. 

(b) One longitudinal test with the 

Hybrid II ATD, deformed floor, with 10 

degrees yaw, and with all lateral 

structural supports (armrests/walls). 

Pass/fail injury assessments: The HIC; 

upper torso restraint load, restraint 

system retention, and pelvic 

acceleration. 

(c) Vertical (14 G’s) test is to be 

conducted with modified Hybrid II 

ATDs with existing pass/fail criteria.

Note: It must be demonstrated that seats 

installed on plinths or pallets meet all 

applicable requirements. Compliance with 

the guidance contained in FAA Policy 

Memorandum PS–ANM–100–2000–00123, 

dated February 2, 2000, titled ‘‘Guidance for 

Demonstrating Compliance with Seat 

Dynamic Testing for Plinths and Pallets’’ will 

be acceptable to the FAA.

Issued in Renton, Washington, on August 

9, 2005. 

Ali Bahrami, 

Manager, Transport Airplane Directorate, 

Aircraft Certification Service.

[FR Doc. 05–16745 Filed 8–22–05; 8:45 am] 

BILLING CODE 4910–13–P

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration 

14 CFR Part 25

[Docket No. NM307; Special Conditions No. 

25–296–SC] 

Special Conditions: Embraer Model 

ERJ 190 Series Airplanes; Sudden 

Engine Stoppage, Interaction of 

Systems and Structures, Operation 

Without Normal Electrical Power, 

Electronic Flight Control Systems, 

Automatic Takeoff Thrust Control 

System (ATTCS), and Protection From 

Effects of High Intensity Radiated 

Fields (HIRF)

AGENCY: Federal Aviation 

Administration (FAA), DOT.

ACTION: Final special conditions.

49156 Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 

SUMMARY: These special conditions are 

issued for the Embraer Model ERJ 190 

series airplane. This airplane will have 

novel or unusual design features when 

compared to the state of technology 

envisioned in the airworthiness 

standards for transport category 

airplanes. These design features are 

associated with (1) engine size and 

torque load which affect sudden engine 

stoppage, (2) electrical and electronic 

systems which perform critical 

functions, and (3) an Automatic Takeoff 

thrust Control Systems (ATTCS). These 

special conditions also pertain to the 

effects of such novel or unusual design 

features, such as their effects on the 

structural performance of the airplane. 

The applicable airworthiness 

regulations do not contain adequate or 

appropriate safety standards for these 

design features. These special 

conditions contain the additional safety 

standards that the Administrator 

considers necessary to establish a level 

of safety equivalent to that established 

by the existing airworthiness standards.

DATES: Effective August 23, 2005.

FOR FURTHER INFORMATION CONTACT: Tom 

Groves, FAA, International Branch, 

ANM–116, Transport Airplane 

Directorate, Aircraft Certification 

Service, 1601 Lind Avenue SW., 

Renton, Washington 98055–4056; 

telephone (425) 227–1503; facsimile 

(425) 227–1149.

SUPPLEMENTARY INFORMATION: 

Background 

Embraer made the original application 

for certification of the ERJ 190 on May 

20, 1999. The Embraer application 

includes six different models, the initial 

variant being designated as the ERJ 190–

100. The application was submitted 

concurrently with that for the ERJ 170–

100, which received an FAA Type 

Certificate (TC) on February 20, 2004. 

Although the applications were 

submitted as two distinct type 

certificates, the airplanes share the same 

conceptual design and general 

configuration. On July 2, 2003, Embraer 

submitted a request for an extension of 

its original application for the ERJ 190 

series, with a new proposed reference 

date of May 30, 2001, for establishing 

the type certification basis. The FAA 

certification basis was adjusted to reflect 

this new reference date. In addition 

Embraer has elected to voluntarily 

comply with certain 14 CFR part 25 

amendments introduced after the May 

30, 2001 reference date. 

The Embraer ERJ 190–100 is a low 

wing, transport-category aircraft 

powered by two wing-mounted General 

Electric CF34–10E turbofan engines. 

The airplane is a 108 passenger regional 

jet with a maximum take off weight of 

51,800 kilograms (114,200 pounds). The 

maximum operating altitude and speed 

are 41,000 feet and 320 knots calibrated 

air speed (KCAS)/0.82 MACH, 

respectively. 

Type Certification Basis 

Based on the May 30, 2001 reference 

date of application, and under the 

provisions of 14 CFR 21.17, Embraer 

must show that the Model ERJ 190 

airplane meets the applicable provisions 

of 14 CFR part 25, as amended by 

Amendments 25–1 through 25–101. If 

the Administrator finds that the 

applicable airworthiness regulations do 

not contain adequate or appropriate 

safety standards for the Embraer ERJ 

190–100 airplane because of novel or 

unusual design features, special 

conditions are prescribed under the 

provisions of 14 CFR 21.16. 

Embraer has proposed to voluntarily 

adopt several 14 CFR part 25 

amendments that became effective after 

the requested new reference date of May 

30, 2001, specifically Amendment 25–

102, except paragraph 25.981(c); 

Amendments 25–103 through 25–105 in 

their entirety; Amendment 25–107, 

except paragraph 25.735(h); 

Amendment 25–108 through 25–110 in 

their entirety; and Amendments 25–112 

through 25–114 in their entirety. 

In addition to the applicable 

airworthiness regulations and special 

conditions, the Embraer Model ERJ 190 

series airplane must comply with the 

fuel vent and exhaust emission 

requirements of 14 CFR part 34 and the 

noise certification requirements of 14 

CFR part 36, and the FAA must issue a 

finding of regulatory adequacy pursuant 

to section 611 of Public Law 93–574, the 

‘‘Noise Control Act of 1972.’’

Special conditions, as defined in 

§ 11.19, are issued in accordance with 

§ 11.38 and become part of the type 

certification basis in accordance with 

§ 21.17(a)(2). 

Special conditions are initially 

applicable to the model for which they 

are issued. Should the type certificate 

for that model be amended later to 

include any other model that 

incorporates the same novel or unusual 

design features, the special conditions 

would also apply to the other model 

under the provisions of § 21.101. 

Discussion of Novel or Unusual Design 

Features 

The Embraer ERJ 190 series airplanes 

will incorporate a number of novel or 

unusual design features. Because of 

rapid improvements in airplane 

technology, the applicable airworthiness 

regulations do not contain adequate or 

appropriate safety standards for these 

design features. The special conditions 

proposed for the Embraer ERJ 190 series 

airplanes contain the additional safety 

standards that the Administrator 

considers necessary to establish a level 

of safety equivalent to that established 

by the existing airworthiness standards. 

These special conditions are the same as 

those required for the Embraer Model 

ERJ 170. 

The Embraer ERJ 190 series airplanes 

will incorporate the novel or unusual 

design features described below.

Engine Size and Torque Load 

Since 1957, § 25.361(b)(1) has 

required that engine mounts and 

supporting structures must be designed 

to withstand the limit engine torque 

load which is posed by sudden engine 

stoppage due to malfunction or 

structural failure, such as compressor 

jamming. Design torque loads associated 

with typical failure scenarios were 

estimated by the engine manufacturer 

and provided to the airframe 

manufacturer as limit loads. These limit 

loads were considered simple, pure 

static torque loads. However, the size, 

configuration, and failure modes of jet 

engines have changed considerably from 

those envisioned when the engine 

seizure requirement of § 25.361(b) was 

first adopted. Current engines are much 

larger and are now designed with large 

bypass fans capable of producing much 

larger torque, if they become jammed. 

Relative to the engine configurations 

that existed when the rule was 

developed in 1957, the present 

generation of engines is sufficiently 

different and novel to justify issuance of 

special conditions to establish 

appropriate design standards. The latest 

generation of jet engines is capable of 

producing, during failure, transient 

loads that are significantly higher and 

more complex than those produced by 

the generation of engines in existence 

when the current regulation was 

developed. 

In order to maintain the level of safety 

envisioned in § 25.361(b), more 

comprehensive criteria are needed for 

the new generation of high bypass 

engines. The proposed special condition 

would distinguish between the more 

common failure events involving 

transient deceleration conditions with 

temporary loss of thrust capability and 

those rare events resulting from 

structural failures. Associated with 

these events, the proposed criteria 

establish design limit and ultimate load 

conditions. 

Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 49157

Interaction of Systems and Structures 

The Embraer Model 190 series 

airplane has fly-by-wire flight control 

systems and other power-operated 

systems that could affect the structural 

performance of the airplane, either 

directly or as a result of a failure or 

malfunction. These systems can 

alleviate loads in the airframe and, 

when in a failure state, can impose 

loads to the airframe. Currently, part 25 

does not adequately account for the 

direct effects of these systems or for the 

effects of failure of these systems on 

structural performance of the airplane. 

The proposed special conditions 

provide the criteria to be used in 

assessing these effects. 

Electrical and Electronic Systems Which 

Perform Critical Functions 

The Embraer Model 190 series 

airplane will have electrical and 

electronic systems which perform 

critical functions. The electronic flight 

control system installations establish 

the criticality of the electrical power 

generation and distribution systems, 

since the loss of all electrical power may 

be catastrophic to the airplane. The 

current airworthiness standards of part 

25 do not contain adequate or 

appropriate standards for the protection 

of the Electronic Flight Control System 

from the adverse effects of operations 

without normal electrical power. 

Accordingly, this system is considered 

to be a novel or unusual design feature, 

and special conditions are proposed to 

retain the level of safety envisioned by 

§ 25.1351(d). 

Section 25.1351(d), ‘‘Operation 

without normal electrical power,’’

requires safe operation in visual flight 

rule (VFR) conditions for at least five 

minutes without normal power. This 

rule was structured around a traditional 

design utilizing mechanical control 

cables for flight control surfaces and the 

pilot controls. Such traditional designs 

enable the flightcrew to maintain 

control of the airplane, while providing 

time to sort out the electrical failure, 

start engines if necessary, and reestablish some of the electrical power 

generation capability. 

The Embraer Model 190 series 

airplane, however, will utilize an 

Electronic Flight Control System for the 

pitch and yaw control (elevator, 

stabilizer, and rudder). There is no 

mechanical linkage between the pilot 

controls and these flight control 

surfaces. Pilot control inputs are 

converted to electrical signals, which 

are processed and then transmitted via 

wires to the control surface actuators. At 

the control surface actuators, the 

electrical signals are converted to an 

actuator command to move the control 

surface. 

In order to maintain the same level of 

safety as an airplane with conventional 

flight controls, an airplane with 

electronic flight controls—such as the 

Embraer Model 190 series—must not be 

time limited in its operation, including 

being without the normal source of 

electrical power generated by the engine 

or the Auxiliary Power Unit (APU) 

generators. 

Service experience has shown that the 

loss of all electrical power generated by 

the airplane’s engine generators or APU 

is not extremely improbable. Thus, it 

must be demonstrated that the airplane 

can continue safe flight and landing 

(including steering and braking on 

ground for airplanes using steer/brakeby-wire) after total loss of normal 

electrical power with the use of its 

emergency electrical power systems. 

These emergency electrical power 

systems must be able to power loads 

that are essential for continued safe 

flight and landing. 

Electronic Flight Control System 

In airplanes with Electronic Flight 

Control Systems, there may not always 

be a direct correlation between pilot 

control position and the associated 

airplane control surface position. Under 

certain circumstances, a commanded 

maneuver that does not require a large 

control input may require a large control 

surface movement, possibly encroaching 

on a control surface or actuation system 

limit without the flightcrew’s 

knowledge. This situation can arise in 

either manually piloted or autopilot 

flight and may be further exacerbated on 

airplanes where the pilot controls are 

not back-driven during autopilot system 

operation. Unless the flightcrew is made 

aware of excessive deflection or 

impending control surface limiting, 

control of the airplane by the pilot or 

autoflight system may be inadvertently 

continued so as to cause loss of control 

of the airplane or other unsafe 

characteristics of stability or 

performance. 

Given these possibilities, a special 

condition for Embraer Model ERJ 190 

series airplanes addresses control 

surface position awareness. This special 

condition requires that suitable display 

or annunciation of flight control 

position be provided to the flightcrew 

when near full surface authority (not 

crew-commanded) is being used, unless 

other existing indications are found 

adequate or sufficient to prompt any 

required crew actions. Suitability of 

such a display or annunciation must 

take into account that some piloted 

maneuvers may demand the airplane’s 

maximum performance capability, 

possibly associated with a full control 

surface deflection. Therefore, simple 

display systems—that would function in 

both intended and unexpected controllimiting situations—must be properly 

balanced between providing needed 

crew awareness and minimizing 

nuisance alerts. 

Automatic Takeoff Thrust Control 

System 

The Embraer Model ERJ 190 series 

airplane will incorporate an Automatic 

Takeoff Thrust Control System (ATTCS) 

in the engine’s Full Authority Digital 

Electronic Control (FADEC) system 

architecture. The manufacturer 

requested that the FAA issue special 

conditions to allow performance credit 

to be taken for use of this function 

during go-around to show compliance 

with the requirement of § 25.121(d) 

regarding the approach climb gradient. 

Section 25.904 and Appendix I refer 

to operation of ATTCS only during 

takeoff. Model ERJ 190 series airplanes 

have this feature for go-around also. The 

ATTCS will automatically increase 

thrust to the maximum go-around thrust 

available under the ambient conditions 

in the following circumstances: 

• If an engine failure occurs during an 

all-engines-operating go-around, or 

• If an engine has failed or been shut 

down earlier in the flight. 

This maximum go-around thrust is 

the same as that used to show 

compliance with the approach-climbgradient requirement of § 25.121(d). If 

the ATTCS is not operating, selection of 

go-around thrust will result in a lower 

thrust level. 

The part 25 standards for ATTCS, 

contained in § 25.904 [Automatic takeoff 

thrust control system (ATTCS) and 

Appendix I], specifically restrict 

performance credit for ATTCS to 

takeoff. Expanding the scope of the 

standards to include other phases of 

flight, such as go-around, was 

considered when the standards were 

issued but was not accepted because of 

the effect on the flightcrew’s workload. 

As stated in the preamble to 

Amendment 25–62:

In regard to ATTCS credit for 

approach climb and go-around 

maneuvers, current regulations preclude 

a higher thrust for the approach climb 

[§ 25.121(d)] than for the landing climb 

[§ 25.119]. The workload required for 

the flightcrew to monitor and select 

from multiple in-flight thrust settings in 

the event of an engine failure during a 

critical point in the approach, landing, 

or go-around operations is excessive. 

Therefore, the FAA does not agree that 

49158 Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 

the scope of the amendment should be 

changed to include the use of ATTCS 

for anything except the takeoff phase. 

(Refer to 52 FR 43153, November 9, 

1987.) 

The ATTCS incorporated on Embraer 

Model ERJ 190 series airplanes allows 

the pilot to use the same power setting 

procedure during a go-around, 

regardless of whether or not an engine 

fails. In either case, the pilot obtains goaround power by moving the throttles 

into the forward (takeoff/go-around) 

throttle detent. Since the ATTCS is 

permanently armed for the go-around 

phase, it will function automatically 

following an engine failure and advance 

the remaining engine to the ATTCS 

thrust level. This design adequately 

addresses the concerns about pilot 

workload which were discussed in the 

preamble to Amendment 25–62. 

The system design allows the pilot to 

enable or disable the ATTCS function 

for takeoff. If the pilot enables ATTCS, 

a white ‘‘ATTCS’’ icon will be displayed 

on the Engine Indication and Crew 

Alerting System (EICAS) beneath the 

thrust mode indication on the display. 

This white icon indicates to the pilot 

that the ATTCS function is enabled. 

When the throttle lever is put in the TO/

GA (takeoff/go-around) detent position, 

the white icon turns green, indicating to 

the pilot that the ATTCS is armed. If the 

pilot disables the ATTCS function for 

takeoff, no indication appears on the 

EICAS. 

Regardless of whether the ATTCS is 

enabled for takeoff, it is automatically 

enabled when the airplane reaches the 

end of the take-off phase (that is, the 

thrust lever is below the TO/GA 

position and the altitude is greater than 

1,700 feet above the ground, 5 minutes 

have elapsed since lift-off, or the 

airplane speed is greater than 140 

knots). 

During climb, cruise, and descent, 

when the throttle is not in the TO/GA 

position, the ATTCS indication is 

inhibited. During descent and approach 

to land, until the thrust management 

system go-around mode is enabled—

either by crew action or automatically 

when the landing gear are down and 

locked and flaps are extended—the 

ATTCS indication remains inhibited. 

When the go-around thrust mode is 

enabled, unless the ATTCS system has 

failed, the white ‘‘ATTCS’’ icon will 

again be shown on the EICAS, 

indicating to the pilot that the system is 

enabled and in an operative condition 

in the event a go-around is necessary. If 

the thrust lever is subsequently placed 

in the TO/GA position, the ATTCS icon 

turns green, indicating that the system 

is armed and ready to operate. 

If an engine fails during the go-around 

or during a one-engine-inoperative goaround in which an engine had been 

shut down or otherwise made 

inoperative earlier in the flight, the 

EICAS indication will be GA RSV (goaround reserve) when the thrust levers 

are placed in the TO/GA position. The 

GA RSV indication means that the 

maximum go-around thrust under the 

ambient conditions has been 

commanded. 

The propulsive thrust used to 

determine compliance with the 

approach climb requirements of 

§ 25.121(d) is limited to the lesser of (i) 

the thrust provided by the ATTCS 

system, or (ii) 111 percent of the thrust 

resulting from the initial thrust setting 

with the ATTCS system failing to 

perform its uptrim function and without 

action by the crew to reset thrust. This 

requirement limits the adverse 

performance effects of a failure of the 

ATTCS and ensures adequate allengines-operating go-around 

performance. 

These special conditions require a 

showing of compliance with the 

provisions of § 25.904 and Appendix I 

applicable to the approach climb and 

go-around maneuvers. 

The definition of a critical time 

interval for the approach climb case is 

of primary importance. During this time, 

it must be extremely improbable to 

violate a flight path derived from the 

gradient requirement of § 25.121(d). 

That gradient requirement implies a 

minimum one-engine-inoperative flight 

path with the airplane in the approach 

configuration. The engine may have 

been inoperative before initiating the goaround, or it may become inoperative 

during the go-around. The definition of 

the critical time interval must consider 

both possibilities. 

Protection From Effects of HIRF 

As noted earlier, Embraer Model ERJ 

190 series airplanes will include an 

Electronic Flight Control System as well 

as advanced avionics for the display and 

control of critical airplane functions. 

These systems may be vulnerable to 

high-intensity radiated fields (HIRF) 

external to the airplane. The current 

airworthiness standards of part 25 do 

not contain adequate or appropriate 

safety standards that address the 

protection of this equipment from the 

adverse effects of HIRF. Accordingly, 

these systems are considered to be novel 

or unusual design features. 

There is no specific regulation that 

addresses protection requirements for 

electrical and electronic systems from 

HIRF. Increased power levels from 

ground-based radio transmitters and the 

growing use of sensitive avionics/

electronics and electrical systems to 

command and control airplanes have 

made it necessary to provide adequate 

protection. 

To ensure that a level of safety is 

achieved that is equivalent to that 

intended by the applicable regulations, 

special conditions are needed for the 

Embraer Model ERJ 190 series airplanes. 

These special conditions require that 

avionics/electronics and electrical 

systems that perform critical functions 

be designed and installed to preclude 

component damage and interruption of 

function due to both the direct and 

indirect effects of HIRF. 

With the trend toward increased 

power levels from ground-based 

transmitters and the advent of space and 

satellite communications coupled with 

electronic command and control of the 

airplane, the immunity of critical 

avionics/electronics and electrical 

systems to HIRF must be established.

It is not possible to precisely define 

the HIRF to which the airplane will be 

exposed in service. There is also 

uncertainty concerning the effectiveness 

of airframe shielding for HIRF. 

Furthermore, coupling of 

electromagnetic energy to cockpitinstalled equipment through the cockpit 

window apertures is undefined. Based 

on surveys and analysis of existing HIRF 

emitters, an adequate level of protection 

exists when compliance with the HIRF 

protection special condition is shown in 

accordance with either paragraph 1 or 2 

below: 

1. A minimum threat of 100 volts rms 

(root-mean-square) per meter electric 

field strength from 10 KHz to 18 GHz. 

a. The threat must be applied to the 

system elements and their associated 

wiring harnesses without the benefit of 

airframe shielding. 

b. Demonstration of this level of 

protection is established through system 

tests and analysis. 

2. A threat external to the airframe of 

the field strengths indicated in the table 

below for the frequency ranges 

indicated. Both peak and average field 

strength components from the table are 

to be demonstrated.

Frequency 

Field strength

(volts per meter) 

Peak Average 

10 kHz–100 kHz ....... 50 50

100 kHz–500 kHz ..... 50 50

500 kHz–2 MHz ........ 50 50

2 MHz–30 MHz ......... 100 100

30 MHz–70 MHz ....... 50 50

70 MHz–100 MHz ..... 50 50

100 MHz–200 MHz ... 100 100

200 MHz–400 MHz ... 100 100

Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 49159

Frequency 

Field strength

(volts per meter) 

Peak Average 

400 MHz–700 MHz ... 700 50

700 MHz–1 GHz ....... 700 100

1 GHz–2 GHz ........... 2000 200

2 GHz–4 GHz ........... 3000 200

4 GHz–6 GHz ........... 3000 200

6 GHz–8 GHz ........... 1000 200

8 GHz–12 GHz ......... 3000 300

12 GHz–18 GHz ....... 2000 200

18 GHz–40 GHz ....... 600 200

The field strengths are expressed in 

terms of peak of the root-mean-square 

(rms) over the complete modulation 

period. 

The threat levels identified above are 

the result of an FAA review of existing 

studies on the subject of HIRF, in light 

of the ongoing work of the 

Electromagnetic Effects Harmonization 

Working Group of the Aviation 

Rulemaking Advisory Committee. 

Discussion of Comments 

Notice of Proposed Special 

Conditions No. 25–05–05–SC for the 

Embraer Model ERJ 190 series airplane 

was published in the Federal Register

dated May 25, 2005 (70 FR 30020). Two 

comments indicating minor errors in the 

proposed special conditions were 

received from Embraer. 

One comment points out an error in 

the table on page 30023 of the Notice of 

Proposed Special Conditions. The 

average level for 1 GHz–2 GHz is shown 

as 2000 rather than as 200 volts per 

meter. The FAA has determined that the 

correct value should be 200 volts per 

meter and, accordingly, has corrected 

the table in these final special 

conditions. 

The second comment indicates an 

error on page 30025. The first sentence 

of Paragraph 2.c.(1)(i) says ‘‘For static 

strength substantiation, these loads 

multiplied by an appropriate factor of 

safety that is related to the probability 

of occurrence of the failure of the 

ultimate loads to be considered for 

design.’’

That sentence should say ‘‘For static 

strength substantiation, these loads 

multiplied by an appropriate factor of 

safety that is related to the probability 

of occurrence of the failure are ultimate 

loads to be considered for design.’’ The 

FAA has determined that the proposed 

wording was incorrect and has corrected 

it in these final special conditions. (We 

also corrected a typographical error in 

the following sentence by removing the 

letter ‘‘I.’’) 

Applicability 

As discussed above, these special 

conditions are applicable to the Embraer 

ERJ 190 series airplane. Should Embraer 

apply at a later date for a change to the 

type certificate to include another 

model incorporating the same novel or 

unusual design features, these special 

conditions would apply to that model as 

well under the provisions of § 21.101. 

Conclusion 

This action affects only certain novel 

or unusual design features of the 

Embraer ERJ 190 series airplane. This is 

not a rule of general applicability.

List of Subjects in 14 CFR Part 25

Aircraft, Aviation safety, Reporting 

and recordkeeping requirements.

The authority citation for these 

special conditions is as follows:

Authority: 49 U.S.C. 106(g), 40113, 44701, 

44702, 44704. 

The Special Conditions 

Accordingly, pursuant to the 

authority delegated to me by the 

Administrator, the Federal Aviation 

Administration (FAA) issues the 

following special conditions as part of 

the type certification basis for the 

Embraer ERJ 190 series airplane. 

Sudden Engine Stoppage 

In lieu of compliance with § 25.361(b) 

the following special condition applies: 

1. For turbine engine installations, the 

engine mounts, pylons and adjacent 

supporting airframe structure must be 

designed to withstand 1g level flight 

loads acting simultaneously with the 

maximum limit torque loads imposed 

by each of the following: 

a. Sudden engine deceleration due to 

a malfunction which could result in a 

temporary loss of power or thrust; and 

b. The maximum acceleration of the 

engine. 

2. For auxiliary power unit 

installations, the power unit mounts 

and adjacent supporting airframe 

structure must be designed to withstand 

1g level flight loads acting 

simultaneously with the maximum limit 

torque loads imposed by each of the 

following: 

a. Sudden auxiliary power unit 

deceleration due to malfunction or 

structural failure; and 

b. The maximum acceleration of the 

power unit. 

3. For engine supporting structures, 

an ultimate loading condition must be 

considered that combines 1g flight loads 

with the transient dynamic loads 

resulting from: 

a. The loss of any fan, compressor, or 

turbine blade; and 

b. Separately, where applicable to a 

specific engine design, any other engine 

structural failure that results in higher 

loads. 

4. The ultimate loads developed from 

the conditions specified in paragraphs 

3.a. and 3.b. above are to be multiplied 

by a factor of 1.0 when applied to 

engine mounts and pylons and 

multiplied by a factor of 1.25 when 

applied to adjacent supporting airframe 

structure. 

Interaction of Systems and Structures 

In addition to the requirements of part 

25, subparts C and D, the following 

special condition applies: 

1. General. For airplanes equipped 

with systems that affect structural 

performance, either directly or as a 

result of a failure or malfunction, the 

influence of these systems and their 

failure conditions must be taken into 

account when showing compliance with 

the requirements of 14 CFR part 25, 

subparts C and D. The following criteria 

must be used to evaluate the structural 

performance of airplanes equipped with 

flight control systems, autopilots, 

stability augmentation systems, load 

alleviation systems, ‘‘flutter’’ control 

systems, and fuel management systems. 

If these criteria are used for other 

systems, it may be necessary to adapt 

the criteria to the specific system. 

a. The criteria defined herein address 

only the direct structural consequences 

of the system responses and 

performances and cannot be considered 

in isolation but should be included in 

the overall safety evaluation of the 

airplane. These criteria may in some 

instances duplicate standards already 

established for this evaluation. These 

criteria are applicable only to structures 

whose failure could prevent continued 

safe flight and landing. Specific criteria 

that define acceptable limits on 

handling characteristics or stability 

requirements when operating in the 

system degraded or inoperative mode 

are not provided in this special 

condition. 

b. Depending upon the specific 

characteristics of the airplane, 

additional studies that go beyond the 

criteria provided in this special 

condition may be required in order to 

demonstrate the capability of the 

airplane to meet other realistic 

conditions, such as alternative gust or 

maneuver descriptions for an airplane 

equipped with a load alleviation system. 

c. The following definitions are 

applicable to this special condition: 

Structural performance: Capability of 

the airplane to meet the structural 

requirements of 14 CFR part 25. 

Flight limitations: Limitations that 

can be applied to the airplane flight 

conditions following an in-flight 

49160 Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 

occurrence and that are included in the 

flight manual (e.g., speed limitations, 

avoidance of severe weather conditions, 

etc.). 

Operational limitations: Limitations, 

including flight limitations, that can be 

applied to the airplane operating 

conditions before dispatch (e.g., fuel 

and payload limitations). 

Probabilistic terms: The probabilistic 

terms (probable, improbable, extremely 

improbable) used in this special 

condition are the same as those used in 

14 CFR 25.1309. 

Failure condition: The term failure 

condition is the same as that used in 14 

CFR 25.1309; however, this special 

condition applies only to system failure 

conditions that affect the structural 

performance of the airplane (e.g., failure 

conditions that induce loads, lower 

flutter margins, or change the response 

of the airplane to inputs, such as gusts 

or pilot actions). 

2. Effects of Systems on Structures.

a. General. The following criteria will 

be used in determining the influence of 

a system and its failure conditions on 

the airplane structure. 

b. System fully operative. With the 

system fully operative, the following 

apply: 

(1) Limit loads must be derived in all 

normal operating configurations of the 

system from all the limit conditions 

specified in 14 CFR part 25, Subpart C, 

taking into account any special behavior 

of such a system or associated functions 

or any effect on the structural 

performance of the airplane that may 

occur up to the limit loads. In 

particular, any significant nonlinearity 

(rate of displacement of control surface, 

thresholds, or any other system 

nonlinearities) must be accounted for in 

a realistic or conservative way when 

deriving limit loads from limit 

conditions.

(2) The airplane must meet the 

strength requirements of 14 CFR part 25 

(static strength, residual strength) using 

the specified factors to derive ultimate 

loads from the limit loads defined 

above. The effect of nonlinearities must 

be investigated beyond limit conditions 

to ensure the behavior of the system 

presents no anomaly compared to the 

behavior below limit conditions. 

However, conditions beyond limit 

conditions need not be considered when 

it can be shown that the airplane has 

design features that will not allow it to 

exceed those limit conditions. 

(3) The airplane must meet the 

aeroelastic stability requirements of 14 

CFR 25.629. 

c. System in the failure condition. For 

any system failure condition not shown 

to be extremely improbable, the 

following apply: 

(1) At the time of occurrence. Starting 

from l-g level flight conditions, a 

realistic scenario, including pilot 

corrective actions, must be established 

to determine the loads occurring at the 

time of failure and immediately after 

failure. 

(i) For static strength substantiation, 

these loads multiplied by an appropriate 

factor of safety that is related to the 

probability of occurrence of the failure 

are ultimate loads to be considered for 

design. The factor of safety (FS) is 

defined in figure 1.

(ii) For residual strength 

substantiation, the airplane must be able 

to withstand two-thirds of the ultimate 

loads defined in subparagraph (c)(1)(i). 

(iii) Freedom from aeroelastic 

instability must be shown up to the 

speeds defined in 14 CFR 25.629(b)(2). 

For failure conditions that result in 

speed increases beyond Vc/Mc, freedom 

from aeroelastic instability must be 

shown to increased speeds, so that the 

margins intended by 14 CFR 

25.629(b)(2) are maintained. 

(iv) Failures of the system that result 

in forced structural vibrations 

(oscillatory failures) must not produce 

loads that could result in detrimental 

deformation of primary structure. 

(2) For the continuation of the flight. 

For the airplane, in the system-failed 

state and considering any appropriate 

reconfiguration and flight limitations, 

the following apply: 

(i) The loads derived from the 

following conditions at speeds up to Vc

or the speed limitation prescribed for 

the remainder of the flight must be 

determined: 

(A) The limit symmetrical 

maneuvering conditions specified in 14 

CFR 25.331 and 25.345. 

(B) The limit gust and turbulent 

conditions specified in 14 CFR 25.341 

and 25.345. 

(C) The limit rolling conditions 

specified in 14 CFR 25.349 and the limit 

unsymmetrical conditions specified in 

14 CFR 25.367 and 25.427(b) and (c). 

(D) The limit yaw maneuvering 

conditions specified in 14 CFR 25.351. 

(E) The limit ground loading 

conditions specified in 14 CFR 25.473 

and 25.491. 

(ii) For static strength substantiation, 

each part of the structure must be able 

to withstand the loads specified in 

paragraph (2)(i) above multiplied by a 

factor of safety depending on the 

probability of being in this failure state. 

The factor of safety is defined in figure 

2.


ER23AU05.023</GPH>
Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 49161

Qj = (Tj)(Pj) where: 

Tj = Average time spent in failure 

condition j (in hours) 

Pj = Probability of occurrence of failure 

mode j (per hour)

Note: If Pj is greater than 10–3 per flight 

hour, then a 1.5 factor of safety must be 

applied to all limit load conditions specified 

in 14 CFR 25, Subpart C.

(iii) For residual strength 

substantiation, the airplane must be able 

to withstand two-thirds of the ultimate 

loads defined in paragraph (c)(2)(ii) 

above. 

(iv) If the loads induced by the failure 

condition have a significant effect on 

fatigue or damage tolerance then their 

effects must be taken into account. 

(v) Freedom from aeroelastic 

instability must be shown up to a speed 

determined from figure 3. Flutter 

clearance speeds V′ and V′′ may be 

based on the speed limitation specified 

for the remainder of the flight using the 

margins defined by 14 CFR 25.629(b).

V′ = Clearance speed as defined by 14 

CFR 25.629(b)(2) 

V′′ = Clearance speed as defined by 14 

CFR 25.629(b)(1) 

Qj = (Tj)(Pj) where: 

Tj = Average time spent in failure 

condition j (in hours) 

Pj = Probability of occurrence of failure 

mode j (per hour)

Note: If Pj is greater than 10–3 per flight 

hour, then the flutter clearance speed must 

not be less than V′′.

(vi) Freedom from aeroelastic 

instability must also be shown up to V′

in figure 3 above for any probable 

system failure condition combined with 

any damage required or selected for 

investigation by 14 CFR 25.571(b). 

(3) Consideration of certain failure 

conditions may be required by other 

sections of 14 CFR 25, regardless of 

calculated system reliability. Where 

analysis shows the probability of these 

failure conditions to be less than 10¥9, 

criteria other than those specified in this 

paragraph may be used for structural 

substantiation to show continued safe 

flight and landing. 

d. Warning considerations. For system 

failure detection and warning, the 

following apply:

(1) The system must be checked for 

failure conditions, not extremely 

improbable, that degrade the structural 

capability below the level required by 

14 CFR part 25 or significantly reduce 

the reliability of the remaining system. 

The flight crew must be made aware of 

these failures before flight. Certain 

elements of the control system, such as 

mechanical and hydraulic components, 

may use special periodic inspections, 

and electronic components may use 

daily checks in lieu of warning systems 

to achieve the objective of this 

requirement. These certification 

maintenance requirements must be 

limited to component failures that are 

not readily detectable by normal 

warning systems and where service 

history shows that inspections will 

provide an adequate level of safety. 

(2) The existence of any failure 

condition not extremely improbable 

during flight—that could significantly 

affect the structural capability of the 

airplane and for which the associated 

reduction in airworthiness can be 

minimized by suitable flight 

limitations—must be signaled to the 


ER23AU05.024</GPH> ER23AU05.025</GPH>
49162 Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 

flight crew. For example, failure 

conditions that result in a factor of 

safety between the airplane strength and 

the loads of 14 CFR part 25, subpart C 

below 1.25 or flutter margins below V’’

must be signaled to the crew during 

flight. 

e. Dispatch with known failure 

conditions. If the airplane is to be 

dispatched in a known system failure 

condition that affects structural 

performance or affects the reliability of 

the remaining system to maintain 

structural performance, then the 

provisions of this special condition 

must be met for the dispatched 

condition and for subsequent failures. 

Flight limitations and expected 

operational limitations may be taken 

into account in establishing Qj as the 

combined probability of being in the 

dispatched failure condition and the 

subsequent failure condition for the 

safety margins in figures 2 and 3. These 

limitations must be such that the 

probability of being in this combined 

failure state and then subsequently 

encountering limit load conditions is 

extremely improbable. No reduction in 

these safety margins is allowed if the 

subsequent system failure rate is greater 

than 10–3 per flight hour. 

Operation Without Normal Electrical 

Power 

In lieu of compliance with 14 CFR 

25.1351(d), the following special 

condition applies: 

It must be demonstrated by test or by 

a combination of test and analysis that 

the airplane can continue safe flight and 

landing with inoperative normal engine 

and APU generator electrical power (in 

other words without electrical power 

from any source, except the battery and 

any other standby electrical sources). 

The airplane operation should be 

considered at the critical phase of flight 

and include the ability to restart the 

engines and maintain flight for the 

maximum diversion time capability 

being certified. 

Electronic Flight Control System 

In addition to compliance with 

§§ 25.143, 25.671 and 25.672, when a 

flight condition exists where, without 

being commanded by the crew, control 

surfaces are coming so close to their 

limits that return to the normal flight 

envelope and (or) continuation of safe 

flight requires a specific crew action, a 

suitable flight control position 

annunciation shall be provided to the 

crew, unless other existing indications 

are found adequate or sufficient to 

prompt that action.

Note: The term suitable also indicates an 

appropriate balance between nuisance and 

necessary operation.

Automatic Takeoff Thrust Control 

System (ATTCS) 

To use the thrust provided by the 

ATTCS to determine the approach climb 

performance limitations, the Embraer 

Model ERJ 190 series airplane must 

comply with the requirements of 

§ 25.904 and Appendix I, including the 

following requirements pertaining to the 

go-around phase of flight: 

1. Definitions 

a. TOGA—(Take Off/Go-Around).

Throttle lever in takeoff or go-around 

position. 

b. Automatic Takeoff Thrust Control 

System—(ATTCS). The Embraer Model 

ERJ–190 series ATTCS is defined as the 

entire automatic system available in 

takeoff when selected by the pilot and 

always in go-around mode, including all 

devices, both mechanical and electrical, 

that sense engine failure, transmit 

signals, and actuate fuel controls or 

power levers or increase engine power 

by other means on operating engines to 

achieve scheduled thrust or power 

increases and to furnish cockpit 

information on system operation. 

c. Critical Time Interval. The 

definition of the Critical Time Interval 

in Appendix I, § I25.2(b) is expanded to 

include the following: 

(1) When conducting an approach for 

landing using ATTCS, the critical time 

interval is defined as 120 seconds. A 

shorter time interval may be used if 

justified by a rational analysis. An 

accepted analysis that has been used on 

past aircraft certification programs is as 

follows: 

(i) The critical time interval begins at 

a point on a 2.5 degree approach glide 

path from which, assuming a 

simultaneous engine and ATTCS 

failure, the resulting approach climb 

flight path intersects a flight path 

originating at a later point on the same 

approach path corresponding to the part 

25 one-engine-inoperative approach 

climb gradient. The period of time from 

the point of simultaneous engine and 

ATTCS failure to the intersection of 

these flight paths must be no shorter 

than the time interval used in evaluating 

the critical time interval for takeoff, 

beginning from the point of 

simultaneous engine and ATTCS failure 

and ending upon reaching a height of 

400 feet. 

(ii) The critical time interval ends at 

the point on a minimum performance, 

all-engines-operating go-around flight 

path from which, assuming a 

simultaneous engine and ATTCS 

failure, the resulting minimum 

approach climb flight path intersects a 

flight path corresponding to the part 25 

minimum one-engine-inoperative 

approach-climb-gradient. The allengines-operating go-around flight path 

and the part 25 one-engine-inoperative, 

approach-climb-gradient flight path 

originate from a common point on a 2.5 

degree approach path. The period of 

time from the point of simultaneous 

engine and ATTCS failure to the 

intersection of these flight paths must be 

no shorter than the time interval used in 

evaluating the critical time interval for 

the takeoff, beginning from the point of 

simultaneous engine and ATTCS failure 

and ending upon reaching a height of 

400 feet. 

(2) The critical time interval must be 

determined at the altitude resulting in 

the longest critical time interval for 

which one-engine-inoperative approach 

climb performance data are presented in 

the Airplane Flight Manual (AFM). 

(3) The critical time interval is 

illustrated in the following figure:

Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 49163

The engine and ATTCS failed time 

interval must be no shorter than the 

time interval from the point of 

simultaneous engine and ATTCS failure 

to a height of 400 feet used to comply 

with I25.2(b) for ATTCS use during 

takeoff. 

2. Performance and System Reliability 

Requirements 

The applicant must comply with the 

following performance and ATTCS 

reliability requirements: 

a. An ATTCS failure or combination 

of failures in the ATTCS during the 

critical time interval: 

(1) Shall not prevent the insertion of 

the maximum approved go-around 

thrust or power or must be shown to be 

an improbable event.

(2) Shall not result in a significant 

loss or reduction in thrust or power or 

must be shown to be an extremely 

improbable event. 

b. The concurrent existence of an 

ATTCS failure and an engine failure 

during the critical time interval must be 

shown to be extremely improbable. 

c. All applicable performance 

requirements of part 25 must be met 

with an engine failure occurring at the 

most critical point during go-around 

with the ATTCS system functioning. 

d. The probability analysis must 

include consideration of ATTCS failure 

occurring after the time at which the 

flightcrew last verifies that the ATTCS 

is in a condition to operate until the 

beginning of the critical time interval. 

e. The propulsive thrust obtained 

from the operating engine after failure of 

the critical engine during a go-around 

used to show compliance with the oneengine-inoperative climb requirements 

of § 25.121(d) may not be greater than 

the lesser of: 

(i) The actual propulsive thrust 

resulting from the initial setting of 

power or thrust controls with the 

ATTCS functioning; or 

(ii) 111 percent of the propulsive 

thrust resulting from the initial setting 

of power or thrust controls with the 

ATTCS failing to reset thrust or power 

and without any action by the crew to 

reset thrust or power. 

3. Thrust Setting 

a. The initial go-around thrust setting 

on each engine at the beginning of the 

go-around phase may not be less than 

any of the following: 

(1) That required to permit normal 

operation of all safety-related systems 

and equipment dependent upon engine 

thrust or power lever position; or 

(2) That shown to be free of hazardous 

engine response characteristics when 

thrust or power is advanced from the 

initial go-around position to the 

maximum approved power setting. 

b. For approval of an ATTCS for goaround, the thrust setting procedure 

must be the same for go-arounds 

initiated with all engines operating as 

for go-arounds initiated with one engine 

inoperative. 

4. Powerplant Controls 

a. In addition to the requirements of 

§ 25.1141, no single failure or 

malfunction or probable combination 

thereof of the ATTCS, including 

associated systems, may cause the 

failure of any powerplant function 

necessary for safety. 

b. The ATTCS must be designed to 

accomplish the following: 

(1) Apply thrust or power on the 

operating engine(s), following any single 

engine failure during go around, to 

achieve the maximum approved goaround thrust without exceeding the 

engine operating limits; 

(2) Permit manual decrease or 

increase in thrust or power up to the 

maximum go-around thrust approved 

for the airplane under existing 

conditions through the use of the power 

lever. For airplanes equipped with 

limiters that automatically prevent the 

engine operating limits from being 

exceeded under existing ambient 

conditions, other means may be used to 

increase the thrust in the event of an 

ATTCS failure, provided that the means 

meet the following criteria: 

• Are located on or forward of the 

power levers; 

• Are easily identified and operated 

under all operating conditions by a 

single action of either pilot with the 

hand that is normally used to actuate 

the power levers, and 

• Meet the requirements of § 25.777 

(a), (b), and (c); 

(3) Provide a means for the flightcrew 

to verify before beginning an approach 

for landing that the ATTCS is in a 

condition to operate (unless it can be 

demonstrated that an ATTCS failure 

combined with an engine failure during 

an entire flight is extremely 

improbable); and 

(4) Provide a means for the flightcrew 

to deactivate the automatic function. 

This means must be designed to prevent 

inadvertent deactivation. 

5. Powerplant Instruments 

In addition to the requirements of 

§ 25.1305, the following requirements 

must be met: 

a. A means must be provided to 

indicate when the ATTCS is in the 

armed or ready condition; and 

b. If the inherent flight characteristics 

of the airplane do not provide adequate 


ER23AU05.026</GPH>
49164 Federal Register / Vol. 70, No. 162 / Tuesday, August 23, 2005 / Rules and Regulations 

warning that an engine has failed, a 

warning system that is independent of 

the ATTCS must be provided to give the 

pilot a clear warning of any engine 

failure during go-around. 

Protection From Effects of HIRF 

Each electrical and electronic system 

that performs critical functions must be 

designed and installed to ensure that the 

operation and operational capability of 

these systems to perform critical 

functions are not adversely affected 

when the airplane is exposed to highintensity radiated fields external to the 

airplane. 

For the purpose of this special 

condition, the following definition 

applies: 

Critical Functions: Functions whose 

failure would contribute to or cause a 

failure condition that would prevent the 

continued safe flight and landing of the 

airplane.

Issued in Renton, Washington, on August 

12, 2005. 

Ali Bahrami, 

Manager, Transport Airplane Directorate, 

Aircraft Certification Service.

[FR Doc. 05–16728 Filed 8–22–05; 8:45 am] 

BILLING CODE 4910–13–P

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration 

14 CFR Part 39

[Docket No. FAA–2005–22145; Directorate 

Identifier 2005–NM–148–AD; Amendment 

39–14223; AD 2005–17–12] 

RIN 2120–AA64

Airworthiness Directives; Bombardier 

Model CL–600–2B19 (Regional Jet 

Series 100 & 440) Airplanes

AGENCY: Federal Aviation 

Administration (FAA), Department of 

Transportation (DOT).

ACTION: Final rule; request for 

comments. 

SUMMARY: The FAA is adopting a new 

airworthiness directive (AD) for certain 

Bombardier Model CL–600–2B19 

(Regional Jet Series 100 & 440) 

airplanes. This AD requires inspecting 

to identify the wing anti-ice ducts 

(piccolo tubes) in the wing leading edge. 

For airplanes with affected piccolo 

tubes, this AD requires revising the 

airplane flight manual (AFM) to 

introduce new procedures for operation 

in icing conditions. The optional 

implementation of repetitive 

inspections for cracks of affected 

piccolo tubes, and corrective actions if 

necessary, terminates the operational 

limitations. The optional installation of 

certain new piccolo tubes terminates 

both the AFM revision and the 

inspections. This AD was prompted by 

reports of failed piccolo tubes. We are 

issuing this AD to prevent cracked 

piccolo tubes, which could result in air 

leakage, a possible adverse effect on the 

anti-ice air distribution pattern and antiice capability without annunciation to 

the flight crew, and consequent reduced 

controllability of the airplane.

DATES: This AD becomes effective 

September 7, 2005. 

The Director of the Federal Register 

approved the incorporation by reference 

of certain publications listed in the AD 

as of September 7, 2005. 

We must receive comments on this 

AD by October 24, 2005.

ADDRESSES: Use one of the following 

addresses to submit comments on this 

AD. 

• DOT Docket Web Site: Go to

http://dms.dot.gov and follow the 

instructions for sending your comments 

electronically. 

• Government-wide Rulemaking Web 

Site: Go to http://www.regulations.gov

and follow the instructions for sending 

your comments electronically. 

• Mail: Docket Management Facility; 

U.S. Department of Transportation, 400 

Seventh Street SW., Nassif Building, 

Room PL–401, Washington, DC 20590. 

• Fax: (202) 493–2251. 

• Hand Delivery: Room PL–401 on 

the plaza level of the Nassif Building, 

400 Seventh Street SW., Washington, 

DC, between 9 a.m. and 5 p.m., Monday 

through Friday, except Federal holidays. 

For service information identified in 

this AD, contact Bombardier, Inc., 

Canadair, Aerospace Group, P.O. Box 

6087, Station Centre-ville, Montreal, 

Quebec H3C 3G9, Canada.

FOR FURTHER INFORMATION CONTACT: Dan 

Parrillo, Aerospace Engineer, Systems 

and Flight Test Branch, ANE–172, FAA, 

New York Aircraft Certification Office, 

1600 Stewart Avenue, suite 410, 

Westbury, New York 11590; telephone 

(516) 228–7305; fax (516) 794–5531.

SUPPLEMENTARY INFORMATION: 

Discussion 

Transport Canada Civil Aviation 

(TCCA), which is the airworthiness 

authority for Canada, notified us that an 

unsafe condition may exist on certain 

Bombardier Model CL–600–2B19 

(Regional Jet Series 100 & 440) 

airplanes. TCCA advises that it has 

received reports of failed wing anti-ice 

ducts (piccolo tubes) located in the wing 

leading edge. De-icing capability was 

degraded on the wing that had the 

piccolo tube damage. Upon 

investigation, it has been determined 

that piccolo tubes manufactured since 

June 2000 are susceptible to cracking 

due to the process used to drill the air 

distribution holes. Such cracking may 

cause air leakage, a possible adverse 

effect on the anti-ice air distribution 

pattern and anti-ice capability without 

annunciation to the flight crew, and 

consequent reduced controllability of 

the airplane. 

Relevant Service Information 

Bombardier has issued Canadair 

Temporary Revision (TR) RJ/155, dated 

July 5, 2005, to the Canadair Regional 

Jet Airplane Flight Manual (AFM), CSP 

A–012. The TR introduces new 

procedures for operation in icing 

conditions. The TR revises the 

Operating Limitations and Abnormal 

Procedures sections of AFM CSP A–012 

to include new procedures for operation 

in icing conditions. 

Accomplishing the actions specified 

in the TR is intended to adequately 

address the unsafe condition. TCCA 

mandated the TR and issued Canadian 

airworthiness directive CF–2005–26, 

dated July 11, 2005, to ensure the 

continued airworthiness of these 

airplanes in Canada. 

Bombardier has also issued Service 

Bulletin 601R–30–029, Revision A, 

dated July 7, 2005. The service bulletin 

describes procedures for: 

• Repetitively inspecting, using 

fluorescent dye penetrant methods, the 

piccolo tubes to detect cracks. 

• Replacing cracked piccolo tubes 

with acceptable parts, or reinstalling 

cracked piccolo tubes under certain 

conditions. 

• Reporting the inspection results to 

the manufacturer. 

FAA’s Determination and Requirements 

of the Proposed AD 

This airplane model is manufactured 

in Canada and is type certificated for 

operation in the United States under the 

provisions of section 21.29 of the 

Federal Aviation Regulations (14 CFR 

21.29) and the applicable bilateral 

airworthiness agreement. Pursuant to 

this bilateral airworthiness agreement, 

TCCA has kept the FAA informed of the 

situation described above. We have 

examined TCCA’s findings, evaluated 

all pertinent information, and 

determined that we need to issue an AD 

for airplanes of this type design that are 

certificated for operation in the United 

States. 

Therefore, we are issuing this AD, 

which requires identifying the part and 

serial numbers of the piccolo tubes 

installed on the airplane. For airplanes