1.0 Missions *

2.0 Global Air Traffic Management (GATM) *

3.0 FMS *

4.0 Autothrottle *

5.0 Flight Deck *

6.0 Station Keeping Equipment (SKE) *

7.0 RADAR Beacon Characteristics *

8.0 Turbulence *

9.0 Air Data System *

10.0 Air Vehicle Electrical System *

11.0 Operational Flight Software *

12.0 Software Verification and Validation *

1. Missions
1. Mission Scenarios

The C-130 provides rapid transport of personnel or cargo for delivery by parachute to a designated drop zone, or by landing at austere locations within the theater of operations. Both methods of delivery require a high degree of precision in both day/night and VMC/IMC. In addition to its theater airlift role, the C-130 has served as a launch platform for remotely piloted vehicles and as a recovery platform for both personnel and data packages. It routinely penetrates hurricanes, provides combat communications links, facilitates rescues on land or at sea, services our remote stations at the North and South Poles, refuels aircraft, broadcasts radio and television messages when the mission requires, and is used to deliver ordnance and fight forest fires. In addition, C-130s also augment strategic airlift forces, and support humanitarian, peacekeeping, and disaster relief operations when needed.

1. SOF Mission Description

1. Current Tactics

The concept of Detection Avoidance Navigation/Threat Avoidance Navigation (DAN/TAN) is used to avoid detection and engagement by hostile radio frequency band, infrared, and optical threats. Historically, SOF aircrews have relied on detailed mission planning, intelligence and low altitude Terrain Following/Terrain Avoidance (TF/TA) flight to reduce the risk of detection and subsequent threat engagement. On currently configured SOF aircraft, aircrews employ conventional TF/TA radar and radar altimeter as primary sensors to enable low altitude terrain following flight. Night Vision Goggles (NVGs) may also be used to supplement these systems where forward visibility and speed permit. Installed electronic warfare (EW) systems coupled with evasive tactics are used to disengage from or defeat threats in the event detection avoidance fails.

2. The Tactical Reality

Given the common use of radar and radar altimeter by penetrating aircraft, hostile air defense forces responded by fielding simple passive detection and radar/communication jamming systems. The success of early passive detection systems has led to the fielding of more sophisticated passive detection naval and land based air defense systems including surveillance aircraft and man-portable surface-to-air missile teams. Modern passive detectors now have the ability to detect low flying aircraft at tactically significant ranges well beyond the line of sight (BLOS) over the horizon. As a result of passive detection, aircraft radars and other radio frequency emitters have become an aid to hostile active Air Defense (AD) forces by providing sufficient warning time to alert and coordinate the use of a wide range of threat systems. Passive detection and the ensuing threat of lethal engagement by hostile forces has forced aircrews to fly at lower altitudes including Nap Of the Earth (NOE). Flight at extremely low altitudes increases the hazard potential of encountering power lines, wires, and towers. Another factor affecting the current tactical situation is the deployment of modern mobile AD systems that have the ability to move rapidly from one location to another and quickly establish lethal combat capability. Mission duration has also experienced an increase as a result of the greater ranges needed for new SOF missions. A requirement has been identified for a material solution that includes updated avionics and EW equipment to combat the present and anticipated threat to SOF aircraft during the accomplishment of low altitude terrain following flight.

3. Identification of the Requirement

1. Global Air Traffic Management (GATM)

Commonality of GATM equipment across AMC weapon systems is desired to reduce the overall AMC support structure, particularly for enroute locations and forward-deployed units. GATM interface with navigation, surveillance, and communication equipment not planned for replacement is imperative.

 

 

2. FMS
1. Navigation Modes

The FMS shall use sensors as shown below:

INU1 shall be the primary INU for the Pilot's integrated Kalman Filter Navigation Solution (KFNS) and INU2 shall be the primary INU for the Copilot's integrated KFNS. If in the event power is removed from the primary INU or the primary INU fails, the user shall be advised.

A user commanded change of the primary INU shall require user verification. This requirement does not prohibit a time constraint for how fast the user can switch back-and-forth between INUs.

1. Kalman Filter Navigation Solution (KFNS)

Each KFNS mode shall have four submodes. The submode shall correspond with the sensors used to navigate in the KFNS mode. The submode shall change automatically according to the functional status of the sensors. The KFNS submodes are listed below in the order of navigation precision, starting with the most accurate submode:

G-I	The G-I submode shall be active when both the INS and GPS data are valid.
GPS	The GPS submode shall be active when the GPS data are valid and INS data are invalid.
D-I	The D-I submode shall be active when the GPS data are invalid and the INS and Doppler radar data are valid.
INS	The INS submode shall be active when neither the GPS nor the Doppler radar data are valid, but the data from at least one INS is valid.

The typical startup sequence on the ground shall follow a predictable upgrade of the KFNS submodes. The KFNS mode corresponds with a single navigation solution, and only one submode may be active at a given time, for a given computer.

The KFNS shall statistically characterize the performance of the GPS, INS, and DVS. The characterized behavior of the INS and Doppler radar, shall be used to calibrate the D-I Kalman filter. If the GPS performance degradation is momentary, the KFNS submode shall return automatically to G-I.

There shall be two other submodes that engage during airborne alignment of the PRIMARY INU. Since these submodes shall be submodes of the KFNS submode, they shall actually be subsubmodes of the INTEGRATED KFNS mode.

GPS-aided In-Air Alignment shall be referred to as GIAA. This shall be submode of the G-I submode.

Doppler radar-aided In-Air Alignment shall be referred to as DIAA. This shall be a submode of the D-I submode.

After an airborne alignment is commanded, the Kalman filter shall operate as an Extended Kalman filter, which means the error estimates shall be feedback to the INU for correcting the INU-computed inertial navigation solution.

2. Independent GPS Mode

The INDEPENDENT GPS solution shall be computed using position from the GPS receiver. Because the GPS position sample rate is only 1 Hz, GPS velocity shall be used to extrapolate position simulate a 4 Hz rate to smooth the display and MCDU indications.

3. Independent Inertial Mode

The INS shall calculate the navigation solution independent of the FMS. Operation in the Inertial-only submode of the INTEGRATED KFNS mode shall be the same as the INDEPENDENT INERTIAL mode except for the following:

The INDEPENDENT INERTIAL solution shall always be calculated when the INU is operative. The Inertial-only submode shall be active only when the Doppler radar and the GPS receiver fail.

In the INDEPENDENT INERTIAL mode, position shall be initialized during airborne alignment. In the Inertial-only submode, position shall be reinitialized to the position of the previously active submode.

In the INDEPENDENT INERTIAL mode, TACAN blending may not be engaged. In the Inertial-only submode, TACAN Blending may be engaged.

The INDEPENDENT INERTIAL navigator may be updated only when the INU is in NAV mode (and not in AA mode). (The FMS navigator in the Inertial-only submode may always be updated, independent of the INU mode.) If the FMS navigator is updated before the INU AA mode is engaged, the position correction shall be feedback to the INU and applied to the INDEPENDENT INERTIAL navigation solution when the AA mode is engaged. Therefore, the INDEPENDENT INERTIAL navigator shall be updated indirectly. If, however, the FMS is in the Inertial-only submode and it is updated while the INU is in AA mode (which means the Doppler radar failed after a DIAA started) the position correction shall not be feedback to the INU, and thus the position update shall be observed for the FMS only.

4. Kalman Filter

A Kalman filter shall provide the integrated navigation function in the G-I, GIAA, D-I, and DIAA submodes of the INTEGRATED KFNS mode. A GPS-Doppler radar Kalman filter shall determine the Doppler radar error characteristics.

The G-I Kalman filter shall operate open loop after a gyrocompass alignment. The G-I Kalman filter shall operate closed loop after an airborne alignment. When the GPS is unavailable, the last GPS-derived estimate of the INU state vector shall be propagated. The G-I Kalman filter shall improve upon the raw GPS navigation by propagating the INU state vector through intermittent periods of GPS performance degradation. The last GPS-derived estimate of the INU state vector shall become the current estimate for correcting the INU.

Kalman filter mode transitions and initializations shall be performed automatically.

The GPS-aided inertial navigator (the G-I Kalman filter) shall not degrade the precision of the existing Doppler radar-aided inertial navigator (the D-I Kalman filter). The D-I Kalman filter shall not be adversely affected by the G-I Kalman filter.

The INU and GPS data shall be synchronized. The Kalman filter shall be initialized in the event the PRIMARY INU changes.

5. Position Updates

The FMS shall provide the following methods of position update according to the sensor used to establish the position fix: TACAN position update, Distance Measuring Equipment (DME)/DME position update, VOR/DME position update, VOR/VOR position, Visual position update, Radar position update, GPS position update, Infrared position update, Shutdown update, and Altitude updates.

The FMS positions updates shall be implemented as follows:

The INS1 and INS2 navigators shall be capable of independent position updates. Updates shall be classified as permanent or temporary. Each class of update is discussed separately in the following paragraphs. Since the GPS is a position sensing system, permanent position update of the GPS-based navigators shall be prohibited. Therefore, neither the INDEPENDENT GPS navigator nor the INTEGRATED KFNS navigator operating in the G-I and GPS-only submodes may be permanently updated. For an airdrop or nonprecision instrument approach procedure, where guidance relative to an unknown or uncertain position may be required, the GPS-based navigators may be fixed temporarily.

1. Permanent Position Updates

Permanent position updates shall provide the following functionality:

Permanent position updates shall reset the angular position coordinates (geodetic latitude and longitude) of the FMS navigation solutions. Update of the vertical position coordinate is a separate operation. Permanent Position Update function shall compare a measure of true position with the FMS estimate of position at a time and place marked by the user.

Permanent position updates shall measure the reference point position relative to the airplane and the FMS estimate of the reference point position relative to the airplane then display the distance between the true and estimated position as the navigation error. Permanent position updates shall allow the user to accept the correction to the FMS navigation solution after displaying the error. The corrections for each FMS solution may be accepted individually.

Permanent position updates shall prohibit position update of the GPS-based navigation solutions (since it is more probable that an update shall increase the error rather than reduce the error).

Permanent position updates shall allow the user to compare the different sensor combinations to observe the relative accuracy of the various types of update.

2. Radar Position Updates

Radar Position updates shall provide the following functionality:

Radar Position updates will operate according to the general FMS permanent position update definition using measurements from the radar. Radar Position updates will be most precise when the cursor coincides with the target return. Any combination of cursor control operations and radar range scales may be used to superimpose the cursor on the target, but the accuracy shall improve as the range scale decreases.

3. Sensor Position Updates

Several variants of the C-130 aircraft fleet have one or more sensors installed on the aircraft that are used to support the aircraft mission. The sensors include television sensors, infrared sensors, radar sensors, and laser sensors. The sensors are capable of providing accurate line-of-sight bearings from the aircraft to a point on the terrain. Some sensors also provide range form the aircraft to a point on the terrain. When the sensors are used to take a fix on terrain points with known locations, the sensor readings shall be made available to the FMS to make permanent position updates. The sensor information generally consists of terrain point bearing (azimuth and elevation) relative to the aircraft plus control and status data. The sensors generally can operate in an independent and slaved mode to facilitate cross check capability. The FMS shall provide the capability to direct any aircraft sensor capable of slaved operation in order to facilitate the location of a specific terrain point.

The interface to aircraft sensor systems will operate according to the general FMS permanent position update definition using measurements from the sensor.

The interface to aircraft sensor systems will allow user input to define the update point in the sensor image as either the reticle center or (if applicable) the track point of the Automatic Video Tracker (AVT). If the update point is the reticle center, the update shall be most precise when the terrain point image appears directly under the reticle in the center of the Field of View (FOV). If the update point is the track point, the update shall be defined according to the area the AVT is tracking.

The interface to aircraft sensor systems will allow the user to employ any combination of applicable manual controller operations and FOV display modes (wide or narrow) to superimpose the reticle center on the terrain point. The interface to aircraft sensor systems will allow the user to engage the AVT.

4. Shutdown Updates

The user shall be able to update the RAW INU position prior to shutdown of the FMS by the use of surveyed coordinates or GPS position if GPS position is using the P Codes with a FOM of 1.

5. Temporary Updates

Temporary position updates shall provide the following functionality:

A temporary position fixing capability shall be provided to guide the airplane relative to a sensor aiming point. This operation shall be referred to as Hot Cursor. Temporary position update shall be provided using measurements from the radar. The term Hot Cursor shall be used to describe the operating state where the position fix defined by the radar cursor is applied directly to the airplane position, without acceptance by the user. Hence, the flight director shall respond directly to the cursor movement.

The system shall provide Hot Cursor steering for the airdrop and ARA procedure. Hot Cursor operation shall continue until the end of the procedure, when flight plan steering shall resume without the temporary position fixing. If (while the cursor is hot) the sensor selected to provide Hot Cursor steering (the radar or the IDS) fails or if the vertical sensor fails, the airplane guidance shall be according to the last valid position fix; that is, the temporary update shall freeze. Hot Cursor steering shall be provided according to the frozen position fix until the airdrop or approach procedure is complete. If the user selects a different sensor to accommodate the failure, and the newly-selected sensor is operative, or if the original sensor becomes operative again, then the temporary fix shall unfreeze and Hot cursor steering shall resume.

For both permanent and temporary position fixing, the selection of an inoperative sensor shall be prohibited once the user has selected a different sensor to accommodate the failure. If all of the alternate sensors have failed and the user attempts to select an alternate sensor, then NO ACCESS shall be displayed at the user interface to indicate the other sensors are inoperative. Position deltas for temporary position update shall be computed and displayed the same as permanent position update. Therefore, the user shall have the option to mark and accept the temporary position fix at a specific instant of time (that is, make the update permanent).

6. Mark Positions

The markpoint (or Mark Position) function shall provide the means for a user to acquire navigation data when the airplane flies over a landmark. The system time, geodetic latitude, and longitude where the mark occurs shall be acquired, and a second user action shall store the markpoint. The user shall be able to define a markpoint via sensor cursor position.

6. Altitude Computation
1. System Altitude

The source of SYSTEM ALTITUDE may be barometric orthometric altitude, baro-inertial orthometric altitude, or GPS orthometric altitude. If the user-selected sensor for SYSTEM ALTITUDE becomes invalid, the alternate sensor shall be selected automatically, and the user shall be advised of the failure and the corresponding change in the source of SYSTEM ALTITUDE. If the user attempts to select a sensor that is invalid, the user shall be advised that the selection is prohibited.

2. GPS Altitude Invalid Conditions

1. User Input for SYSTEM ALTITUDE Calculation

A user input shall define the vertical position/range sensor for the functions that require a measure of aircraft altitude or ground clearance, or target elevation or target clearance.

2. Radar Altimeter Ground Clearance Measurement

Dual Radar Altimeters shall be integrated into the AMP system and provide ground clearance information from 0 to 50,000 ft. Dual Radar Altimeters shall provide a visual and aural low altitude warning indication if the measured altitude drops below a manually set limit. Their accuracy shall be ± 2% from 0 – 5,000 ft. and ± 1% above 5,000 ft.

3. Altitude Updates

1. Typical airdrop geometries

Typical airdrop geometries are shown in Figure 1 and Figure 2.



Figure 1. Airdrop Geometry

 

 



Figure 2. IMC Airdrop Profile

2. Rendezvous Guidance and Steering

The FMS shall have the capability of performing a rendezvous function. The rendezvous function shall predict the intercept point with a moving target and provide guidance to the rendezvous. The time-tagged position and velocity of the target shall be entered on the MCDU. The position, velocity, and information provided by TCAS shall be used to determine the intercept point and course. Guidance and steering shall be provided to the intercept point.

3. Orbital guidance

Orbits shall be specified by the following parameters:

Orbit point, selectable either as a preloaded waypoint or via map coordinates (e.g. latitude, longitude, elevation)

Orbit altitude, either AGL or MSL

Orbit radius

Left or right hand orbit

Cone angle, defined as the angle between an aircraft body fixed vector (e.g. gun barrel) and the vertical. When cone angle is used, either altitude or radius may be specified, but not both.

Orbit rate, either as airspeed or orbit period.

This function shall calculate an orbit with the selected parameters, and generate cues to enter and maintain that orbit. Reasonableness checks shall be performed, and the crew alerted if an inconsistent set of parameters is entered, the resulting orbit is beyond the capabilities of the aircraft, the resulting orbit violates operational limits (including, but not limited to, bank, airspeed, height above terrain, g’s), or if the resulting orbit requires performance exceeding the capabilities of the autopilot thus preventing coupled operation.

Results of the calculation shall be displayed to the flight crew via MFD. This display shall include the maximum and minimum bank angle required, airspeed required, as well as all the above parameters. A plan view of the orbit overlaid onto a navigation chart shall also be provided.

Engagement of orbit steering to the MFD and HUD shall require a positive crew action, allowing the crew to follow other steering cues while previewing the orbit. Coupling of orbit guidance to the autopilot shall also require a positive crew action. Orbit guidance shall be discontinued when so commanded by the crew or when higher priority guidance is activated. In that case, the crew shall be alerted that guidance mode has changed. Calculation of the orbit shall compensate for estimated wind.

4. Lateral and Vertical Navigation
1. Lateral Leg Transitions

Same as ARINC Characteristic 702A paragraph 4.3.3.1.2 with the following additions:

The FMS shall be capable of sequencing waypoints manually or automatically, depending on the user's preference. The GOTO waypoint shall be sequenced automatically when automatic transitions are enabled and the transition point of the active leg has been passed. The user shall be able to manually command a transition to a new leg by selecting either a new GOTO waypoint or a new active leg except for waypoints associated with landing and drop zones.

2. Airdrop Guidance and Steering Functions
1. Airdrop With Temporary Position Update

Temporary position fixing shall be activated during an airdrop procedure in order to steer the airplane relative to a sensor aiming point. Refer to "Temporary Position Updates". During an airdrop, Hot Cursor steering shall be provided according to a user-defined offset aiming point, a user-selected vertical sensor, a user-selected aiming sensor (for example, slant range or bearing), and the manual trimming commands for the aiming sensor.

The steering signals for the Flight Director shall be based on the position that corresponds with Flight Director mode, but the temporary position corrections (that correspond with the temporary position update) shall be relative to the position of the INTEGRATED KFNS navigator. That is, the corrections shall be determined from the INTEGRATED KFNS navigator, and then applied to the guidance solution selected to operate the Flight Director. If an FMS-independent Flight Director mode is engaged, both the steering signals and the temporary position update shall be based on the position of the INTEGRATED KFNS navigator.

If the user uncouples the lateral sensor from the FMS (for example, the user engages the manual mode of the sensor in order to operate the sensor independent of the FMS), then Hot Cursor steering shall be deactivated and the temporary position correction shall be reset to zero.

The release point shall be a function of the down range and vertical range of the point of impact, the payload ballistics, and the wind. The ground range shall be a function of the position that corresponds with both the Flight Director mode and the user-selected aiming sensor for temporary position fixing. The vertical sensor for determining the airplane position relative to the Altitude Gate shall be the same sensor selected for computing the release point. The troop jump light system and cargo release mechanism shall be activated at the release point.

2. Airdrop Without Temporary Position Update

Lateral guidance shall be provided to the release point in accordance with the flight plan. For the FMS-dependent Flight Director modes, the signals for the Flight Director shall be based on the position that corresponds with the Flight Director mode. If an FMS-independent Flight Director mode is engaged, the steering signals shall be based on the position of the INTEGRATED KFNS navigator.

The release point shall be based on the down range and vertical range of the point of impact, the payload ballistics, and the wind. A single user input shall select a common sensor for SYSTEM ALTITUDE, for computing the release point, and for the Altitude Gate. The troop jump light system and the cargo release mechanism shall be activated at the release point.

3. Airborne Sensor Approach (ASA)

The FMS shall provide lateral and vertical guidance for the ASA function described in the SRD and in the following sections.

1. ASA With Temporary Position Update

Temporary position fixing shall be activated during the approach procedure in order to steer the airplane relative to a sensor aiming point. During ARA, Hot Cursor steering shall be provided according to a user-defined offset aiming point, a user-selected vertical sensor, a user-selected aiming sensor (for example, slant range or bearing), and manual trimming commands for the aiming sensor.

The lateral steering signals for the Flight Director shall be based on the position that corresponds with the Flight Director mode, but the temporary position corrections (that correspond with the temporary position update) shall be relative to the position of the INTEGRATED KFNS navigator. That is, the corrections shall be determined from the INTEGRATED KFNS navigator, and then applied to the guidance solution selected to operate the Flight Director.

If an FMS-independent Flight Director mode is engaged, both the steering signals and the temporary position update shall be based on the position of the INTEGRATED KFNS navigator. If the user uncouples the aiming sensor from the FMS (for example, the user engages the manual mode of the sensor in order to operate the sensor independent of the FMS), then Hot Cursor steering shall be deactivated and the temporary position correction shall reset to zero.

Vertical guidance shall be based on the down range and vertical range of the touchdown point. The ground range shall be a function of both the position that corresponds with the Flight Director mode and the user-selected aiming sensor for temporary position fixing. The vertical sensor for vertical guidance shall be the sensor selected for SYSTEM ALTITUDE. Therefore, a single user input shall define a common sensor for both SYSTEM ALTITUDE and vertical guidance.

During active approach to a predefined LZ waypoint, the FMS shall generate vertical steering indicators that, if followed, shall guide the aircraft to capture the glideslope. Glideslope capture guidance shall be based on operator and sensor inputs to the FMS, and the output shall be the glideslope deviation indicator. Also, guidance information and advisories shall be output to the MFD. The FMS shall perform the glideslope capture function based on operator input. These entries shall include glideslope, LZ elevation, height above the touchdown elevation, distance to the Missed Approach Point, and the Missed Approach Point distance.

The FMS shall also process inputs for altitude (GPS altitude or barometric altitude) or ground clearance (CARA radar altimeter) from the aircraft sensors. When the required entries have been made, the minimum and maximum altitude for glideslope capture shall be computed and displayed.

Upon transition of the IP waypoint enroute to the LZ, the FMS shall display the altitude range for glideslope capture as MIN/MAX ALT. If the aircraft altitude is outside the minimum/maximum glideslope capture altitude range during an approach, the FMS shall display ARA GATE (flashing) in the message window on each MCDU.

2. ASA Without Temporary Position Update

1. Autothrottle
1. Autothrottle Performance Requirements

The AWFCS should include an autothrottle capability that performs the following functions: airspeed control in all phases of flight including SKE (station keeping equipment) formation , airdrop operations, approach, autoland, and go-around, and hold of a manually selected airspeed. All functions shall be available in the range of idle to maximum forward thrust. When the aircraft is stabilized in a climb, cruise, descent, or coordinated turn mode and the throttle is not operating at the minimum or maximum limits, the autothrottle shall maintain the aircraft speed within ± 5 knots of the engaged airspeed under the following conditions:

Airspeed: 1.2 x Vs to 0.64M/318 knots CAS

Altitude: Sea level to 45,000 feet

Gross Weight: 90,000 to 175,000 pounds

The autothrottle function shall control the throttle movement in response to speed and vertical flight path commands. There shall be no undesirable periodic oscillations of the throttle commands. There shall be no transient engagement oscillations. The autothrottle function shall provide fail-passive. When engaged, the automatic flight control function shall provide the autothrottle functions defined in Table 4.1.

Autothrottle shall disconnect when any engine exceeds torque and TIT (turbine inlet temperature) limits as defined in T.O. 1C-130-1. The throttle force (with the autothrottle function not engaged) shall be 6.5 pounds nominal. The autothrottle function shall be operational during stabilized climb, cruise, descent, and coordinated turns. All modes shall be available from idle to maximum forward thrust. The AMP architecture shall allow the pilot to physically overpower the autothrottle function with a nominal force of 16 lbs. per throttle.

Table 4.1. Autothrottle Function (Engaged) Performance Limits

Mode or Submode1	Control or Sensor	Parameter	Limits
Airspeed/Mach Hold (AMAH) Note 2, 5, 6	Autothrottle Control Central Air Data Computer (CADC)	Command Targets Tolerance TIT Control Range 5   TIT Hold Error Torque Mach Engage Range Mach Hold Error KCAS Engage Range 6 KCAS Hold Error Throttle Control Authority Throttle Rate Limit   Wind Shear/ Gust Compensation Settling Time         Residual Oscillations	± 1 KCAS or ± 0.01 M 17 to 106 %, not to exceed 875° C TIT (T56-7) 1010° C TIT (T56-15) ± 2% 1.2 Vstall to 0.64 M ± 0.01 M 100 to 318 KCAS ± 5 KCAS 29° (idle) to 80° (max) ± 7° /sec Throttle Angular Velocity (TAV) Note 7, 8   Following engagement or perturbation of this mode at 2000 fpm or less, the specific ALH accuracy shall be achieved within 30 seconds Period shall not be less than 20 seconds
VNAV/FMS   Submodes: Climb Cruise Descent	Autothrottle Control FMS CADC	TIT Control Range   TIT Hold Error Mach Engage Range Mach Hold Error IAS Engage Range IAS Hold Error Throttle Control Authority Throttle Rate Limit	17 to 106 %, not to exceed 875° C TIT (T56-7) 1010° C TIT (T56-15) ± 2 % 1.2 Vstall to 0.64 M ± 0.01 M 100 to 318 KCAS ± 5 knots 29° (idle) to 80° (max) ± 7° /sec TAV
Take Off /Go Around (TOGA)3	Autothrottle Control Go Around Button	TOGA Command Targets TIT Control Range   TIT Hold Error Throttle Control Authority Throttle Rate Limit	Manual Control 17 to 106 %, not to exceed 875° C TIT (T56-7) 1010° C TIT (T56-15) ± 2 % 29° (idle) to 80° (max) ± 20° /sec TAV
Approach (Autoland)4 (APPR)	Autothrottle Control	TIT Control Range   TIT Hold Error Throttle Control Authority Throttle Rate Limit Flare Limits (Throttle Retard) Throttle Range Limit	17 to 106 %, not to exceed 875° C TIT (T56-7) 1010° C TIT (T56-15) ± 2 % 29° (idle) to 80° (max) ± 7° /sec TAV Active at 50 ft AGL 10% RMS nominal throttle position

NOTES:

1 The disengaged submode shall be available for use, if it is compatible with the other engaged modes.

2 Adjustments allowed up to full authority with airspeed hold command control in increments of ± 1.0 KCAS or ± 0.005 Mach..

3 During go around the autothrottle function shall drive the throttles to the takeoff position within 4 seconds.

4 The autopilot function shall provide the autothrottle function with the command to retard the throttles during the flare maneuver.

5 The autothrottle clutchpack shall be designed to preclude overheat.

6 Engagement of the autothrottle in steady-state conditions when the difference between aircraft speed and selected speed is within 5 knots, shall not cause more than 1.5 degrees of throttle action.

7 The airspeed error shall be held within 2% of clutched-in airspeed in a non-linear wind shear of up to 5 knots per 100 feet with aircraft sink rates up to 1,000 feet per minute.

8 In vertical or longitudinal gusts, the maximum airspeed error shall not exceed 3% of the clutched-in airspeed.

2. Flight Deck

The AMP cockpit arrangement shall be designed using JSSG-2010 as a guide.

3. Station Keeping Equipment (SKE)

As a minimum, the current AN/APN-169C SKE capability shall be retained and integrated.

Replacement of the existing AN/APN-169C SKE system is desired. If replaced, the replacement system shall, as a minimum provide the following capabilities: Allow aircraft to perform precision airdrops, rendezvous, air refueling, and airland missions at night and in all weather conditions to include instrument meteorological conditions (IMC). The system shall allow as few as 2 aircraft and as many as 100 (250 desired) aircraft to maintain formation position/separation at selectable ranges from 500 feet to 100 NMs at all IFR altitudes. The system shall allow multiple formations to interfly and maintain formation position/ separation at selectable distances from 500 feet to 100 NMs at all IFR altitudes. The system shall have all weather intraformation positioning/collision avoidance capability with all similarly equipped aircraft and with all aircraft currently IMC formation equipped. The system shall be compatible with current and future ground based zone marker (ZM) or compatible systems and shall interrogate system within 40 NMs (100 NMs desired). The system shall provide relative position information on all aircraft in the formation, or a subset of selected aircraft (i.e., element or serial) to include distance, bearing, heading, airspeed, and relative altitude. The system shall provide steering commands to correct and maintain formation position settings. The system shall provide visual and aural proximity and collision warnings of similarly equipped aircraft and other aircraft currently IMC formation equipped that infringe on selected range and provide warnings for loss of signal or system degradation. The system shall be capable of being coupled to and interfacing with the autopilot during all phases of SKE operation including airdrop.

4. RADAR Beacon Characteristics

Parameter	SST-181X-E	PPN-19	SMP-1000
Output to Radar (Beacon RF)
Response Freq. (MHz)	8800-9500 (9310 nom.)	9250-9450 (9310 nom.) 16.25± 0.10 GHz	9310
Pulse Width	0.3 ± 0.1 m s	0.35 ± 0.05 m s	0.375 ± 0.025 m s
Pulse Coding	Code Spacing (NM) 1 0 (single pulse) 2 4 3 5 4 6 5 7 6 8 7 9 8 10 9 11 10 12	Code Pulse Spacing (NM) A 1 B 4 C 1, 2 alternating D 1, 4 alternating E 3, 4 alternating F 2, 4 alternating G 0 (single pulse)	SST-181 Codes 1-5 PPN-19 Codes A-G Strobe
Polarization	Horizontal	Horizontal	Horizontal
Input from Radar (Radar RF)
Interrogation Freq.	8800-9500 MHz (9375 nom)	9375 MHz 16-17 GHz	8800 - 9800 MHz
Pulse Width	0.25 – 5.0	I- and J-bands: 0.2 – 3.0 m sec GAR-I band: 0.25 – 3.0 m sec	0.3 m sec min.
Polarization	Horizontal	Horizontal	Horizontal

 

 

5. Turbulence

A further objective is detection of turbulence. The system should measure the spectral width of each weather return and declare turbulence present whenever the one-sigma value (standard deviation) of the spectral width is equal to or greater than 5 m/s. The system should provide data to the display system such that areas of turbulence can be identified using unique colors or techniques to readily alert the operator. The system should be capable of detecting turbulence to a range of 50 NM.

6. Air Data System
1. Pitot – Static System

The total overload wattage of heaters at any time shall not exceed 1200 watts. In still air at an ambient temperature between 20 and 30 degrees F. , the heater shall not exceed 275 watts after 5 minutes of heater operation.

The pitot pressure line which extends aft to the pitot pressure chamber and on to the pressure fitting shall consist of 0.25 inch outside diameter tubing having a wall thickness of 0.022 inch. The fitting shall be in accordance with MIL-F-5509 and MS33656-4. The static pressure line in the tube and extending to the fitting shall consist of 0.25 inch outside diameter tubing having a wall thickness of 0.022 inch. The static pressure line fitting shall be in accordance with MIL-F-5509 and MS33656-6. The static pressure line in the aircraft shall have a 3/8 inch outside diameter with a 0.035 wall thickness for aluminum tubing. The pitot pressure lines in the aircraft shall have a ¼ inch outside diameter with a wall thickness of 0.035 inch for aluminum tubing. Supports for the tubing shall be provided at appropriate lengths. Lines shall be appropriately marked near any fitting to distinctly show which line is static and which is pitot.

 

 

 

7. Air Vehicle Electrical System

ESU provides MIL-STD-704E AC power to the Main and Essential Avionics Buses. The ESU electrical system TCTO 1C-130-1339 is comprised of 2 Bus Switching Systems (BSS), each consisting of a 10 KVA inverter, Rectifier Control Unit, and a Base assembly. The ESU Main and Essential AC Avionics buses are rated at 10KVA 3-phase power per bus. However, each bus is limited to 3.3KVA per phase per bus. ESU also replaces the 250VA and 2500VA inverters with form, fit, and function units. ESU replaces the Frequency Sensor Relay and Voltage Regulators with Generator Control units (GCU) and updates the overhead electrical control panel.

8. Operational Flight Software
1. Application Program Interface (API)

1. Software Verification and Validation

Contractor shall have a disciplined, standardized software verification and validation process. This process includes technical and documentation reviews, quality and configuration audits, software process/product measurements, and software certification (testing). Contractor shall develop a Verification and Validation plan which ensures that the software functionality is correctly implemented and that the customer’s software requirements have been achieved.

1. Software Testing/Certification