The Triple-Spool Core

This article walks the length of the Trent 700 itself: three concentric shafts, the compressors and turbines on them, the combustion system between them, the bearings that locate them, and the two layers of insurance built in for the day a fan blade — or the LP shaft itself — lets go. Everything that attaches the engine to the aircraft (pylon, cowls, accessory gearbox, drains) is the next article; everything the engine does with its own air (variable stators, bleed valves, internal cooling) is article 03. Here we deal with the rotating machine.

1. Why three spools

The AMM opens with the engine's official identity:

"The RB211-TRENT engine is a high by-pass ratio, triple spool turbo-fan. It includes an LP system and a gas generator system. Most of the engine thrust comes from the LP system which is supplied with power from the gas generator."

That sentence contains a more useful decomposition than the familiar "seven systems" list: the engine is really two machines in series. The gas generator (IP compressor + HP compressor + combustor + HP and IP turbines) exists to produce hot, high-energy gas. The LP system (the fan and the 4-stage LP turbine that drives it) exists to convert that energy into thrust. Most of the air the fan moves never enters the core at all — it leaves through the bypass duct, and it carries most of the thrust. The core burns fuel, in the end, to spin a very large propulsor.

Why split the core machinery across three shafts rather than the more common two? Because each section of compression has a different optimum speed. The fan is enormous; its tips go supersonic — and its efficiency collapses — if it turns too fast, so it must turn slowly. The HP compressor handles small, dense, hot air and must turn fast to do useful work per stage. A twin-spool design forces two compromises; a triple-spool design lets the LP, IP and HP rotors each sit at their own best speed, with no mechanical connection between them — only aerodynamic coupling through the gas path. The price is mechanical complexity: three nested shafts and a bearing arrangement that takes a diagram to appreciate (§2.2). The visible cockpit consequence: three spool speeds, N1/N2/N3, where a twin-spool aircraft shows two.

2. Anatomy in one pass

2.1 The longitudinal section

   LP COMPRESSOR     IP COMPRESSOR    HP COMPR.   HP TURB  IP TURB   LP TURBINE
      (fan)             (8 stages)    (6 stages) (1 stage)(1 stage)  (4 stages)
        ▒▒                ▨▨▨▨▨▨▨▨      ██████      █        ▨        ▒▒▒▒
        ▒▒ ─────┐         ▨▨▨▨▨▨▨▨      ██████      █        ▨        ▒▒▒▒
   ◄────▒▒      │  bypass duct (cold stream = most of the thrust) ─────────►
        ▒▒      └─▨▨▨▨▨▨▨▨─██████─[combustor]─█──▨──▒▒▒▒──► common nozzle
        ▒▒    ╔═══════════════════════════════════════════╗
        ▒▒    ║ LP shaft (innermost, full length: fan──LP turbine)
        ▒▒    ║   IP shaft (middle: IP compressor──IP turbine)
        ▒▒    ║     HP drum (outermost, shortest: HP compr.──HP turbine)
        ▒▒    ╚═══════════════════════════════════════════╝
        ▒▒        │
                  ▼  (via internal gearbox → radial/inclined drive shaft)
            ACCESSORY GEARBOX (external, under the fan case → article 02)

Three reading points. First, the shading pairs each compressor with the turbine that drives it — fan with LP turbine, IP with IP, HP with HP — three closed loops that share nothing but gas. Second, note the nesting: the LP shaft is the longest and innermost, running the full length of the engine; the IP shaft sits around it; the HP "shaft" is the shortest of all and is really a drum. That nesting dictates the shaft-failure protection design in §4. Third, the accessory gearbox is not on the engine centreline — it hangs under the fan case and takes its power from the HP spool through an internal gearbox and an inclined drive shaft (§8).

2.2 The bearing arrangement

Per the AMM compressor description:

"The three sets of location bearings, one for each shaft, are contained in the internal gearbox. The roller bearings are contained in three different bearing compartments."

Three details matter operationally. All three location (thrust) bearings live together in the internal gearbox — the axial position of all three shafts is managed by one structural assembly. The HP and IP turbine roller bearings share a single housing, which is why the EGT probes and the turbine-bearing oil services all cluster in the HP/IP turbine case area. And the LP turbine rear bearing carries a spring preload: behind the roller bearing sits a ball-bearing-and-spring set that pushes the LP shaft assembly continuously rearward, limiting axial float on the longest, most slender shaft in the machine.

2.3 The engine's identity card

"Engine Data: Engine Mark: RB211-TRENT 768-60, Bare Engine Take-off Thrust: 67 500 / RB211-TRENT 772-60, Bare Engine Take-off Thrust: 71 100. Direction of Rotation: Counterclockwise when seen from the rear. Pressure Ratio: 37.42:1. Bypass Ratio: 4.66:1. Compressor Stages: 1 LP, 8 IP, 6 HP. Combustion: Annular Combustion Chamber with 24 Fuel Spray Nozzles. Turbine Stages: 1 HP, 1 IP, 4 LP."

The bypass ratio deserves one plain-language sentence: for every unit of air that goes through the core and burns fuel, the fan pushes 4.66 units of cold air that never sees the combustor. "Most of the thrust is in the bypass" is not rhetoric — it is the flow split.

3. The gas path, step by step

The FCOM compresses the whole working cycle into five numbered steps — the pilot-level minimum complete story:

"The engine operates as follows: 1. The LP compressor, referred to as the fan, compresses the air. 2. Then, the air is divided into two flows: ‐ Most of the air flows out of the core engine, and provides most of the engine thrust, ‐ The remaining air enters the core engine. 3. The IP and the HP compressors compress the air that enters the core engine. 4. The fuel is added to and mixed with the compressed air of the core engine. The mixture is ignited in the combustion chamber. 5. The gas that results from combustion drives the HP, the IP, and the LP turbines."

Step 5 hides the explanation for the stage-count asymmetry. The gas hits the HP turbine first, then IP, then LP — so the turbine closest to the combustor receives gas at its highest energy grade. One HP stage is enough to feed six stages of HP compression precisely because it drinks the freshest, hottest, highest-pressure gas. By the time the flow reaches the LP turbine it has surrendered most of its energy, so four stages are needed to wring out enough power for the huge fan. The turbine stage count — 1, 1, 4 — is a direct mirror of the declining energy grade along the gas path.

At the back, the AMM closes the loop:

"As the gas flow comes out of the turbine it mixes with the LP compressor cold flow." — "Behind the LP turbine there is a common nozzle assembly which mixes the cold air and hot gas exhaust flows."

The Trent 700 is a mixed-exhaust engine: hot and cold streams merge in the Common Nozzle Assembly (CNA) and leave through one nozzle. The exhaust description adds two refinements worth keeping: the flows that merge are actually three — fan cold flow, cool bleed-air discharge, and the hot turbine exhaust — and the CNA is a convergent duct that accelerates the mixed flow and contributes additional thrust. In the AMM's own words, "the exit area nozzle controls the engine exit thrust and has a large effect on engine performance" — the nozzle is not a passive outlet but part of the thrust-producing machinery. The mixed-exhaust layout is also part of why reversing only the fan stream is sufficient for the thrust reverser (article 13).

4. N1, N2, N3 — and a trap that must be unlearned deliberately

"The rotation speed of the fan provides the N1 engine parameter. The rotation speed of the IP rotor provides the N2 engine parameter. The rotation speed of the HP rotor provides the N3 engine parameter. The N1 and N3 engine parameters appear on the Engine/Warning Display (E/WD). The N2 engine parameter appears on the ENG SD page."

[!warning]- If your muscle memory says "N2 = the high-pressure spool = what I watch during start" — it is wrong here On a triple-spool engine the numbering simply runs front to back: N2 is the intermediate spool; N3 is the high-pressure spool. The starter cranks N3; ignition and fuel-on timing gates are read against N3; the permanently displayed pair on the E/WD is N1 + N3, while N2 — counter-intuitively the least prominent spool — appears only on the ENG SD page. Every twin-spool habit that keys on "N2" must be consciously remapped to N3 on this engine. The start sequence built on N3 gates is article 12.

The FCOM also assigns the parameters their jobs:

"The FADEC uses: ‐ The N1 engine parameter to compute the applicable engine thrust, ‐ The N1, N2 and N3 engine parameters for engine control and monitoring."

N1 computes thrust; all three monitor the machine. This coexists without contradiction with the fact that the commanded thrust-setting parameter in normal operation is EPR — EPR is the control-loop target, N1 the thrust computation and fallback dimension. The full relationship is article 07.

5. The LP system — fan, containment, and two layers of shaft insurance

The fan itself. One stage, and per the AMM:

"The LP compressor is a one stage rotor with 26 wide-chord type blades which engage in axial dovetail slots. Each blade is held in the disk with two shear keys, and radial movement is prevented by slider assemblies."

Twenty-six wide-chord blades — few and broad — is the combination that buys both aerodynamic efficiency and bird-strike tolerance. Annulus fillers between the blade roots fair the hub line smooth.

Insurance layer one: containment. If a blade does let go, its fragment carries enormous energy, and the case must hold it:

"Wound around the outside of this framework, to include the compressor blade track, is a bandage of kevlar material. The kevlar bandage has great strength and if a LP compressor blade breaks, the kevlar material will contain it."

The case construction is a three-layer division of labour: an aluminium isogrid framework (72 axial ribs plus diagonals) carries stiffness; the kevlar bandage wound over the blade track provides containment; a titanium outer skin provides the fire barrier. Like body armour, the containment layer is not trying to be rigid — it is trying to catch and dissipate.

Insurance layer two: the failsafe shaft. The LP shaft is the longest, most slender shaft in the engine. If it fractured, the fan — suddenly free of the rearward pull that the LP turbine normally transmits — would drive itself forward out of the nacelle. The Trent 700 therefore carries a second shaft inside the first:

"If the compressor shaft breaks, separation of the LP compressor from the engine is prevented by a failsafe shaft. Installed internally, this shaft is kept in position at the rear by a collar and a nut attached to the turbine shaft. At the front, the failsafe shaft is attached to the LP compressor shaft with nuts and bolts."

Note what the failsafe shaft does and does not do: it prevents separation only — it transmits no drive. With the main shaft broken the engine stops producing thrust regardless; the fan is simply held captive instead of becoming a free-flying flywheel. Its control-system partner is the FADEC's LP turbine overspeed protection (the unloaded LP turbine spins up violently the instant the shaft breaks, and the FADEC cuts fuel immediately — article 05, article 28). Mechanical capture plus control-system fuel cut: a two-layer closure for the same failure.

The LP turbine (module 08): four disks bolted into a drum, stage 3/4 blades installed as welded pairs; the stage-3 disk carries both the forward curvic coupling to the LP shaft and the rearward stubshaft into the rear bearing.

6. The IP system — eight stages, and the bearing housing that doubles as a sensor house

"The IP compressor has 8 stages, and the IP turbine has a single stage." — "To optimize and protect the operation of the engine at all power settings and in all flight conditions, the Variable Inlet Guide Vanes (VIGVs) and the Variable Stator Vanes (VSVs) regulate the quantity of air that flows through the IP compressor."

Read that second sentence carefully: on this engine the variable geometry lives on the IP compressor — 58 VIGVs behind the engine-section stators, VSVs on IP stages 1 and 2, driven by unison rings on the titanium half-case — not on the HP compressor. All of the engine's adjustable surge protection acts on the IP section; the mechanism and its control law are article 03.

The Front Bearing Housing (FBH) is an underrated component — not merely a bearing support but the engine's sensor house:

"Installed in the FBH hub are the LP and IP shaft roller bearings and an oil sump. Also, there are the speed probes which measure both LP and IP compressor speeds. Two of the ESS vanes have electrical cables installed internally to transmit signals from the pick-ups. Six more vanes contain tubes to supply oil to and from the roller bearings."

Both the N1 and N2 speed signals originate inside the FBH — phonic wheels on the LP shaft and the IP front stubshaft read by electromagnetic pick-ups — and the structural vanes of the engine section do double duty as conduits: two vanes carry the sensor wiring, six more carry bearing oil, all without disturbing the gas path. The full signal chain is article 14.

The IP turbine (module 05): a single stage of 126 blades on firtree roots, preceded by 26 hollow NGVs of which 13 conceal structural struts that carry the HP/IP bearing housing, while others route oil lines and IP8 cooling air — once again, aerofoils used as plumbing.

7. The HP system — and the geography of EGT

"The HP compressor has 6 stages, and the HP turbine has a single stage."

"The rotor blades installed in stages 1 to 3 are made of titanium and the others, of a heat resistant alloy."

That material change mid-compressor is physical evidence worth pausing on: by stage 4, compression heating (an overall pressure ratio of 37:1 does not come free) has taken the air beyond what titanium tolerates comfortably, so stages 4–6 switch to a heat-resistant alloy. The rear of a high-pressure compressor is far hotter than intuition suggests. At the front of the HP drum, a stubshaft rides in the HP location bearing and drives the internal gearbox — the power take-off for everything in §8.

The HP turbine: a single stage of 92 blades, firtree roots locked in pairs, shrouded tips forming a continuous sealing ring against the case, with seal fins ahead of the disk guiding cooling air into the blade roots. One stage, 92 small blades, the highest speed and highest temperature in the machine — this is the most fragile component of the engine and the true object of every EGT limit.

Where EGT is actually measured — the chapter's first big counter-intuitive:

"Adjacent to the rear flange there is a cooling air manifold, and the location bosses for eleven thermocouples." — "The first stage of NGVs, which are hollow, are installed as 3-vane sets in the HP/IP case. One vane in each of eleven sets contains a thermocouple, and another set includes an overheat detector."

The eleven EGT thermocouples sit in the first-stage nozzle guide vanes of the LP turbine — downstream of both the HP and IP turbines. EGT is therefore the temperature of gas that has already spent most of its energy driving two turbines; the true combustor exit temperature is far higher, but no sensor material would survive there for long. Think of it as a thermometer halfway up the chimney rather than inside the stove: the reading is offset low, but as long as the conversion is stable, watching the chimney controls the stove. This is why the EGT limits (900/920 °C, article 00) look "far below combustion temperature" yet are hard red lines — they are HP-turbine protection limits back-computed through a known relationship. Display and signal chain: articles 14/15; exceedance handling: article 28.

8. Combustion — 24 nozzles, 2 igniters, and why dilution must arrive late

"The combustion chamber burns a mixture of fuel and HP air. The FADEC controls the fuel/air mixture in accordance with the position of the thrust lever and the aircraft operating conditions. The combustion chamber is an annular assembly with fuel nozzles and two igniters. The combustion chamber is between the HP compressor and the HP turbine."

The AMM supplies the numbers and a beautiful piece of reasoning. The annular chamber is built from inner and outer liners; fuel enters through 24 spray nozzles on the burner ring; two igniter plugs reach in through bosses in the outer case, which also carries eight borescope ports. The liners are perforated with holes and dilution chutes, and the placement of those chutes encodes the mechanism:

"Air is added along the chamber, through dilution chutes, after combustion has occurred. This is to cool the gas flow. The air must be added after combustion because if not, the air will cool the flame and combustion will not be complete."

Only a small fraction of the air actually burns with the fuel — that keeps the flame core hot and stable. The rest arrives late and in two roles: film-cooling air hugs the liner walls to protect the liner itself, and dilution air joins downstream to blend the roughly 2 000 °C flame core down to a temperature the turbine can survive. Let the fire burn completely first; add the cold air afterwards. Reverse the order and you get incomplete combustion or a quenched flame. Why there are two igniters, where they sit and what feeds them is article 11; how fuel reaches the 24 nozzles is article 09.

9. The accessory power tap

The FCOM gives the drive list that every pilot should be able to recite:

"The accessory gearbox drives various accessories with mechanical power via the HP shaft for the operation of the engine and the aircraft systems. The accessory gearbox of each engine operates: ‐ The oil feed pump that provides the oil system with oil. ‐ The main engine fuel pump that provides the combustion chamber with fuel. ‐ The engine-driven hydraulic pumps that pressurize the GREEN, the BLUE and the YELLOW hydraulic systems. ‐ The engine-driven generators that are the primary source of electrical power. ‐ The FADEC alternator that provides the FADEC with electrical power. ‐ The pneumatic starter that enables the engine start."

The mechanical route: HP drum front stubshaft → internal gearbox (housed inside the ten structural vanes of the intermediate case) → inclined/radial drive shaft → external accessory gearbox under the fan case. The aircraft's entire electrical generation, two of the three hydraulic systems' engine pumps, and the engine's own oil, fuel and FADEC power all draw from this single HP-shaft tap — the hardware root of the principle from article 00 that an engine failure is never "just thrust". During start the path runs in reverse: the pneumatic starter drives N3 through the same gearbox, which is exactly why a start is monitored on N3 and not N1. The gearbox itself, accessory by accessory, is article 02.

10. Modules and borescope geography

"The basic engine is an assembly of primary units which are identified as modules. These modules can be independently replaced and are specified as follows: module 01: LP compressor rotor / 02: IP compressor / 03: Intermediate case / 04: HP system (this includes the HP compressor, the combustion system and the HP turbine) / 05: IP turbine / 06: External gearbox / 07: LP compressor case / 08: LP turbine."

Note that module 04 packages the HP compressor, the combustor and the HP turbine as one unit — the hottest, fastest-turning heart of the machine is replaced as a block. For a pilot the practical meaning is in maintenance language: "module 04" on a borescope report means the core heart; and the borescope ports scattered along the cases (IP stages 3/5/7, HP stages 1/2/5, eight combustor ports, the turbine stages) exist so that the heart can be inspected without opening the chest. Every overtemperature, stall or overspeed you write up is typically followed by an inspection through exactly these ports.

11. Where this structure meets the failure chapters

This is a pure structure article — it owns no ECAM alert — but the failure spectrum stands on it:

Self-test

[!note]- Q1. What do "1-8-6" and "1-1-4" stand for, and which shaft carries which? Compressor stages LP–IP–HP = 1, 8, 6; turbine stages HP–IP–LP = 1, 1, 4. The fan (1 stage) is driven by the 4-stage LP turbine; the 8-stage IP compressor by the single IP turbine; the 6-stage HP compressor by the single HP turbine. The turbine counts mirror the falling energy grade of the gas as it works rearward.

[!note]- Q2. Which spool does the starter crank, and which two spool speeds live permanently on the E/WD? The starter drives N3 (the HP spool) through the accessory and internal gearboxes. The E/WD permanently shows N1 and N3; N2 appears only on the ENG SD page. Any twin-spool habit keyed to "N2" must be remapped to N3.

[!note]- Q3. A fan blade fails. What contains it — and what holds the fan if the LP shaft fails instead? A broken blade is caught by the kevlar containment bandage wound over the blade track of the isogrid fan case. A broken LP shaft is a different event: the internal failsafe shaft prevents the fan from separating forward, while the FADEC's LP turbine overspeed protection cuts fuel as the unloaded turbine accelerates — mechanical capture plus control-layer fuel cut.

[!note]- Q4. Where exactly do the eleven EGT thermocouples sit, and why does that location matter? In the first-stage LP turbine nozzle guide vanes — downstream of both the HP and IP turbines. EGT is therefore much lower than combustor exit temperature; the limits built on it protect the HP turbine through a stable, known offset. It is a thermometer in the chimney, not in the stove.

[!note]- Q5. Why must dilution air enter the combustor only after combustion? Because added earlier it would chill the flame and leave combustion incomplete. Only a small air fraction burns; film-cooling air protects the liners and dilution air then blends the flame core down to turbine-tolerable temperature.

Key takeaways

References

• FCOM DSC-70 (powerplant description) — five-step working cycle, stage counts, N1/N2/N3 definitions and display split, FADEC parameter usage, combustor description, accessory gearbox drive list.
• AMM 72-00 (engine general, D/O) — engine data table, LP/gas-generator decomposition, module list, exhaust mixing.
• AMM 72-30 (compressors, D/O) — fan blade construction, kevlar containment, failsafe shaft, FBH contents, VIGV/VSV location, HP blade materials, bearing arrangement.
• AMM 72-40 (combustion, D/O) — liners, 24 nozzles, igniter and borescope bosses, late-dilution rationale.
• AMM 72-50 (turbines, D/O) — HP 92 blades, IP 126 blades and strut-bearing NGVs, LP turbine drum, EGT thermocouple bosses, rear-bearing spring preload.
• AMM 78-11 (exhaust, D/O) — CNA three-flow mixing, convergent nozzle and exit-area effect on thrust.
• Integrative synthesis (marked in text): the optimum-speed rationale for three spools; the energy-grade reading of stage counts; the stove/chimney analogy for EGT; the materials-change interpretation.

Independent study material, not an Airbus publication and not endorsed by the manufacturer. Always defer to the current operator FCOM, FCTM, and QRH for operational use.