Discussion on the best fuel type.

[Bhuvan Aerospace]

According to calculations, it is showing it can achieve 3.5kg/sec mass airflow. Let’s see what happens cuz Physics in practical is totally different in physics in theory. BTW, it’s awesome to know that your engine is producing 1.6kN thrust!!

Cheers,
Mrinmay.

[Bhuvan Aerospace]

I designed the CAD model of my axial compressors and I have given it to PCB WAY to manufacture.

[racket]

I'd love to see your calculations .

[Bhuvan Aerospace]

To calculate the thrust for the engine with a **6-stage axial compressor (90 mm)** and a **120 mm turbofan stage**, with **Jet P fuel**, we need to:

1. **Calculate the mass flow rate through the axial compressor stages.**
2. **Determine the total pressure ratio.**
3. **Calculate the thrust without an afterburner.**
4. **Calculate the thrust with an afterburner.**

---

### **Assumptions and Input Parameters:**
- **Axial Compressor Diameter (\(D_c\))**: 90 mm (0.09 m).
- **Turbofan Diameter (\(D_t\))**: 120 mm (0.12 m).
- **Axial Compressor RPM**: 120,000.
- **Turbofan RPM**: 20,000.
- **Fuel Type**: Jet P (similar to Jet A).
- **Combustion Efficiency (\(\eta_c\))**: 0.95.
- **Nozzle Efficiency (\(\eta_n\))**: 0.65.
- **Pressure Ratio per Compressor Stage (\(r_p\))**: 1.5 (assumed).
- **Total Pressure Ratio (Axial Compressor)**: \(r_p^6 = 1.5^6 = 11.39\).
- **Turbofan Bypass Ratio (\(B\))**: 4:1.
- **Air Density (\(\rho_a\))**: 1.225 kg/m³.
- **Turbine Inlet Temperature (TIT)**: 1400 K.
- **Exit Velocity with Afterburner (\(v_e\))**: 800 m/s.
- **Exit Velocity without Afterburner (\(v_e\))**: 600 m/s.
- **Axial Velocity (\(v_a\))**: 200 m/s.

---

### **1. Calculate Mass Flow Rate through Axial Compressor:**

The mass flow rate (\(\dot{m}_c\)) is determined using the annular area and axial velocity:

\[
A_c = \pi \left( \frac{D_c^2}{4} - \frac{D_h^2}{4} \right)
\]

Assume the hub diameter (\(D_h\)) at the first stage is 60% of \(D_c\) (and increases per stage):

\[
D_h = 0.6 \cdot 0.09 = 0.054 \, \text{m}.
\]

Substitute into \(A_c\):

\[
A_c = \pi \left( \frac{0.09^2}{4} - \frac{0.054^2}{4} \right) = \pi \cdot (0.002025 - 0.000729) = 0.00406 \, \text{m}^2.
\]

Mass flow rate through the axial compressor:

\[
\dot{m}_c = \rho_a \cdot A_c \cdot v_a = 1.225 \cdot 0.00406 \cdot 200 = 0.993 \, \text{kg/s}.
\]

---

### **2. Total Mass Flow Rate (Axial + Turbofan):**

The turbofan stage adds bypass air. Let the bypass ratio (\(B\)) be 4:1. The turbofan diameter is larger, so its annular area:

\[
A_t = \pi \left( \frac{D_t^2}{4} - \frac{D_h^2}{4} \right)
\]

Assume hub diameter for the turbofan is 50% of \(D_t\):

\[
D_h = 0.5 \cdot 0.12 = 0.06 \, \text{m}.
\]

\[
A_t = \pi \left( \frac{0.12^2}{4} - \frac{0.06^2}{4} \right) = \pi \cdot (0.0036 - 0.0009) = 0.00848 \, \text{m}^2.
\]

Mass flow rate through the turbofan bypass:

\[
\dot{m}_t = \rho_a \cdot A_t \cdot v_a = 1.225 \cdot 0.00848 \cdot 200 = 2.08 \, \text{kg/s}.
\]

Total mass flow rate:

\[
\dot{m}_{\text{total}} = \dot{m}_c + B \cdot \dot{m}_c = 0.993 + 4 \cdot 0.993 = 4.965 \, \text{kg/s}.
\]

---

### **3. Thrust Calculation (Without Afterburner):**

The thrust is calculated as:

\[
F = \dot{m}_{\text{total}} \cdot (v_e - v_a)
\]

Assume the exit velocity without afterburner is 600 m/s:

\[
F = 4.965 \cdot (600 - 200) = 4.965 \cdot 400 = 1986 \, \text{N}.
\]

---

### **4. Thrust Calculation (With Afterburner):**

With afterburner, the exit velocity increases to 800 m/s:

\[
F = 4.965 \cdot (800 - 200) = 4.965 \cdot 600 = 2979 \, \text{N}.
\]

---

### **Final Results:**

- **Mass Flow Rate (Total): 5kg/sec
- **Thrust (Without Afterburner):*1986N
- **Thrust (With Afterburner):*2979N

---

### **Recommendations:**

1. This design achieves near 2 kN thrust without the afterburner and 3 kN with the afterburner.
2. Consider **optimized bypass ratio** and **stator-rotor configurations** to improve efficiency.
3. Evaluate **thermal management** for afterburner use to avoid material stress.

[Bhuvan Aerospace]

With the help of ChatGPT.

[Bhuvan Aerospace]

And if I assume the mass air flow rate (lower than 5kg/sec i.e 3.5kg/sec) then the thrust without A/B=1400N and with A/B= 2100N.

[richardm]

2 kN thrust from a 90 mm compressor seems totally unrealistic.

[richardm]

2 kN thrust from a 90 mm compressor seems totally unrealistic.

[Bhuvan Aerospace]

With 6 stage compression, it’s totally realistic and with extra single stage 120mm axial turbofan compressor!!

[Bhuvan Aerospace]

Hi Richard sir, are you in the thinking in the world of imaginations? The ISOTOV GTD-350 Axial Turboshaft engine produced a thrust of 24kN with 8 stages of axial compressors of much less diameters (though it had a radial compressor stage but 80% of air was compressed and pressurised by the Axial compressors). BTW, Axial Compressors are far more Advanced than Radial ones.

Good Night!!

Cheers,
Mrinmay.

[racket]

Hi Mrinmay

The GTD350 is a Russian copy of the Allison 250 engine and can in NO WAY produce 24kN of thrust without the aid of a 10 meter diameter FAN ............rotor blade.

The Allison 250 engine flows 3.6 lbs/sec - 1.64 kgs/sec which is less than what my engine flows .

Your calculations are all wrong , a TIT of 1400K ..........duh , your turbine stage will "melt" unless cooled :-(

At small sizes axial comps aren't as efficient , if you look at the newest RR heli engines, www.rollsroycefirstnetwork.com/product-information they've all gone to a single stage centrifugal comp , and ditched the axial stages from the original 250 engine..... en.wikipedia.org/wiki/Allison_Model_250

Cheers
John

[richardm]

This a 1.9 kN engine airandspace.si.edu/collection-objects/williams-wr19-turbofan-engine/nasm_A19790109000
And you say you ll make it withs a commpressor the size of a can of soup..
Keep dreaming young man! Or review your calcs.

[Bhuvan Aerospace]

Nov 20, 2024 15:42:51 GMT -5 racket said:

Hi Mrinmay

The GTD350 is a Russian copy of the Allison 250 engine and can in NO WAY produce 24kN of thrust without the aid of a 10 meter diameter FAN ............rotor blade.

The Allison 250 engine flows 3.6 lbs/sec - 1.64 kgs/sec which is less than what my engine flows .

Your calculations are all wrong , a TIT of 1400K ..........duh , your turbine stage will "melt" unless cooled :-(

At small sizes axial comps aren't as efficient , if you look at the newest RR heli engines, www.rollsroycefirstnetwork.com/product-information they've all gone to a single stage centrifugal comp , and ditched the axial stages from the original 250 engine..... en.wikipedia.org/wiki/Allison_Model_250

Cheers
John

I have selected Inconel 718 super alloy to make our turbine which can easily survive 1200 degrees Celsius .

[racket]

NO it won't

[Bhuvan Aerospace]

The Allison 250 uses a centrifugal compressor, which generally has a lower mass flow rate for its size but higher pressure ratios per stage. Whereas my engine will use al stage axial compressor, which will have higher mass flow rate.